Study of the feasibility of a “Rocket launching Consultancy” through the analysis of the propulsion system’s requirements to reach LEO and MEO orbits with payloads up to 1,000 kg

Author: Robert Arcaleanu

Director: Borja Pedro Borràs Quintanal

Co-Director: David Bermejo Plana

Bachelor’s Thesis

Bachelor’s degree in Aerospace Vehicle Engineering

Escola Superior d’Enginyeries Industrial, Aeroespacial i Audiovisual de Terrassa Universitat Politècnica de Catalunya

June, 30th 2020

Acknowledgement

On the one hand, I would like to thank to my director and co-director, Borja Borràs and David Bermejo, who provided me the opportunity to develop a project related to what I like, rockets. It would not be possible without them, and I show them my sincere gratitude.

On the other hand, I would like to express my gratitude to all my friends that have supported me during the last four years. Last but not least, I would like to thank to my parents, who always have been there for everything I needed.

C

Contents

AcknowledgementC

Contentsi

List of Figuresv

List of Tables vii

Abstract xiii

Aim xv

Scope xvii

Requirements xix

Background xxi

1 Introduction1 1.1 Brief Rockets’ Historical Review...... 2 1.2 Business Description and Feasibility Analysis...... 3 1.2.1 Potential Customer...... 4

2 Market Research5 2.1 Buying a Launch Service...... 5 2.2 Competitors...... 6 2.2.1 NASA Launch Services Program...... 6 2.2.1.1 NASA Venture Class Launch Services...... 7 2.2.2 Spaceflight...... 7 2.2.3 Precious Payload...... 8 2.2.4 Loft Orbital...... 9

i ii CONTENTS

2.2.5 EXOLAUNCH ...... 9 2.2.6 D-Orbit ...... 9 2.2.7 Other Launch Services Providers ...... 9

3 State of the art 11 3.1 Payload ...... 11 3.1.1 Human Spacecraft ...... 11 3.1.2 Cargo Spacecraft ...... 12 3.1.3 Space Probes ...... 13 3.1.4 Satellites...... 13 3.2 Rocket Launch Sites ...... 14 3.2.1 Geographic Considerations...... 15 3.2.1.1 Latitude...... 15 3.2.1.2 Azimuth...... 15 3.2.2 Active Launch Sites...... 16 3.3 Operating Launch Vehicles...... 17 3.3.1 Rocket Features Definition...... 18 3.3.2 Launch vehicles...... 19 3.4 Orbits ...... 21 3.4.1 Orbit Parameters...... 21 3.4.2 Earth orbit classification...... 22

4 Study of Propulsion System’s Requirements 25 4.1 Missions Parameters and Definition...... 25 4.2 Study of the Possibility to Reach the Orbits ...... 30 4.2.1 Possibility to Reach the Polar Orbit...... 32 4.2.2 Possibility to Reach the Sun-Synchronous Orbit ...... 33 4.2.3 Possibility to reach the Inclined Orbit...... 33 4.2.4 Possibility to Reach the Semi-Synchronous Orbit ...... 34 4.2.5 Possibility to Reach the Molniya Orbit...... 34 4.3 Propulsion System’s Parameters Calculation...... 35 4.3.1 Delta V Budget...... 35 4.3.2 Rocket Equations...... 39 4.3.3 Optimal Staging ...... 41 4.3.3.1 Results and Algorithm Validation...... 48

ii CONTENTS ESEIAAT

5 Results Analysis 51 5.1 Results Comparison...... 51 5.1.1 Polar Orbit ...... 52 5.1.2 Sun-Synchronous Orbit...... 52 5.1.3 Inclined Orbit...... 53 5.1.4 Semi-Synchronous Orbit...... 54 5.1.5 Molnyia Orbit...... 55 5.2 Feasibility Study ...... 57 5.2.1 Other Services...... 58

6 Conclusions and Future Work 61

Environmental Impact 62

Budget 64

Planning 66

A Operating Launch Vehicles 73 A.1 Two Stage Launch Vehicles...... 73 A.2 Three Stage Launch Vehicles...... 75 A.3 Four Stage Launch Vehicles ...... 76 A.4 Five Stage Launch Vehicles ...... 77

B Possibility to Reach the Orbits 79 B.1 Polar Orbit ...... 80 B.2 Sun-Synchronous Orbit...... 81 B.3 Inclined Orbit...... 82 B.4 Semi-Synchronous Orbit ...... 83 B.5 Molnyia Orbit...... 84

C Results of Optimal Staging 85 C.1 Polar Orbit ...... 86 C.2 Sun-Synchronous Orbit...... 87 C.3 Inclined Orbit...... 89 C.4 Semi-Synchronous Orbit ...... 90 C.5 Molnyia Orbit...... 92

iii iv CONTENTS

D Optimisation Script 95

Bibliography 99

Declaració d’Honor 103

iv List of Figures

2.1 Spaceflight Launch Vehicle Family[1]...... 8

3.1 Velocity at Earth’s Surface by Latitude [2] ...... 15 3.2 Potential Launch Azimuths from Omelek[3] ...... 17 3.3 Palmachim and Sohae Hypothetical Azimuths . . . . 17 3.4 Serial Staging [4]...... 18 3.5 Parallel Staging [4]...... 18 3.6 Orbital Elements [5] ...... 21 3.7 Earth Orbits Representation...... 23

4.1 Launch Vehicle Performance and Launch Mission. . . . 26 4.2 Topocentric horizon coordinate system [6]. A is the azimuth...... 30 4.3 Launch Azimuth[7]...... 30 4.4 Hohmann Transfer[6]...... 31 4.5 Launch Vehicle Boost Trajectory[8]...... 36 4.6 Typical Atlas V 521 Dynamic Pressure vs Time[9] . . . 36 1 2 4.7 First order approximation ( 2 ρv vs t)...... 36 4.8 Launch System Performance Losses[6] ...... 38 4.9 Definitions of Various Vehicle Masses[10] ...... 40 4.10 Chamber and Nozzle Representation [10] ...... 42 4.11 Altitude performance of RS 27 liquid propellant rocket engine used in early versions of the Delta launch vehi- cle [10]...... 43 4.12 Lift-off Mass for a Two Stages Rocket to Reach 200 km Circular Orbit (i = 51.6o) with 1,000 kg of Payload 47 4.13 Optimal Staging with Graphical Method (Falcon 9) . . 48 4.14 Optimal Staging with Graphical Method (Rockot) . . . 49

v vi LIST OF FIGURES

6.1 Gantt Diagram (1/2)...... 70 6.2 Gantt Diagram (2/2)...... 71

vi List of Tables

2.1 LSP Fleet for Launching Payloads into LEO or beyond [11]..... 7

3.1 Human Spacecraft Specifications...... 12 3.2 Cargo Spacecraft Specifications [12]...... 12 3.3 Space Probes Features [13][14]...... 13 3.4 Satellite Classification according to Mass [15]...... 14 3.5 Active Orbital Launch Sites [2][16] ...... 16 3.6 Launch vehicle’s sites...... 20

4.1 Difference between GLONASS, GPS and Galileo [17] ...... 29 4.2 Delta-V to Reach the Parking Orbit (Polar Orbit) ...... 32 4.3 Delta-V to Reach the Parking Orbit (Sun-Synchronous Orbit) . . . 33 4.4 Delta-V to Reach the Parking Orbit (Inclined Orbit) ...... 33 4.5 Delta-V to Reach the Parking Orbit (Semi-Synchronous Orbit) . . . 34 4.6 Delta-V to Reach the Parking Orbit (Molniya Orbit) ...... 34 4.7 Payload Fairing Dimensions ...... 37 4.8 Space Shuttle Incremental Flight Velocity Breakdown[10] . . . . . 39 4.9 Advantages and Disadvantages of Chemical Propellants[6] . . . . . 41 4.10 Performance of Chemical Propellants[10]. *These are reference values. 42 4.11 First Stage Engines[18]...... 44 4.12 Upper Stages Engines[19]...... 44 4.13 Comparison between the Graphical Method and the Lagrange Mul- tiplier Method (Falcon 9)...... 49 4.14 Comparison between the Graphical Method and the Lagrange Mul- tiplier Method (Rockot) ...... 50 4.15 Comparison between Real Mass and Computed Mass (Minotaur-I) . 50

5.1 Optimal Launch Vehicles to Reach Polar Orbit according to the Mission ...... 52

vii viii LIST OF TABLES

5.2 Optimal Launch Vehicles to Reach Sun-Synchronous Orbit accord- ing to the Mission...... 52 5.3 Optimal Launch Vehicles to Reach Inclined Orbit according to the Mission ...... 53 5.4 Optimal Launch Vehicles to Reach Semi-Synchronous Orbit accord- ing to the Mission...... 54 5.5 Optimal Launch Vehicles to Reach Molnyia Orbit according to the Mission ...... 55 5.6 Missions Classification ...... 56

6.1 Rocket and Aviation Emissions[Metric T onnes] [20]...... 63 6.2 Software Licenses Budget...... 65 6.3 Energy and Equipment Budget ...... 66 6.4 Total Budget ...... 66 6.5 Bachelor’s Thesis Tasks...... 68 6.6 Level of effort to develop each task ...... 69

A.1 Two Stage Launch Vehicles without Boosters Configuration [21] . . 73 A.2 Two Stage Launch Vehicles with Boosters Configuration [21] . . . . 73 A.3 Angara Launch Vehicle Family [21] ...... 74 A.4 Atlas V Launch Vehicle Family [21][9]...... 74 A.5 Delta IV Launch Vehicle Family [21][22] ...... 74 A.6 H-II Launch Vehicle Family [21][19]...... 74 A.7 Three Stage Launch Vehicles without Boosters Configuration [21] . 75 A.8 Three Stage Launch Vehicles with Boosters Configuration [21] . . . 75 A.9 Four Stage Launch Vehicles Without Booster Configuration [21] . . 76 A.10 PSLV Launch Vehicle Family [21][23]...... 76 A.11 Five Stage Launch Vehicles without Booster Configuration [21]... 77

B.1 Calculations of thePossibility to Reach the Polar Orbit ...... 80 B.2 Calculations of the Possibility to Reach the Sun-Synchronous Orbit 81 B.3 Calculations of the Possibility to Reach the Inclined Orbit . . . . . 82 B.4 Calculations of the Possibility to Reach Semi-Synchronous Orbit . 83 B.5 Calculations of the Possibility to Reach the Molniya Orbit . . . . . 84

C.1 Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Polar Orbit...... 86 viii LIST OF TABLES ESEIAAT

C.2 Optimal Three Stage Launch Vehicles according to the Propellant Type to Reach Polar Orbit...... 86 C.3 Optimal Four and Five Stage Launch Vehicles according to the Pro- pellant Type∗ to Reach Polar Orbit...... 87 C.4 Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Sun-Synchronous Orbit...... 87 C.5 Optimal Four and Five Stage Launch Vehicles according to the Pro- pellant Type to Reach Sun-Synchronous Orbit...... 88 C.6 Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Inclined Orbit ...... 89 C.7 Optimal Two Stage Launch Vehicles according to the Propellant Type∗ to Reach Inclined Orbit...... 89 C.8 Optimal Four and Five Stage Launch Vehicles according to the Pro- pellant Type∗ to Reach Inclined Orbit...... 90 C.9 Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Semi-Synchronous Orbit...... 90 C.10 Optimal Three Stage Launch Vehicles according to the Propellant Type∗ to Reach Semi-Synchronous Orbit...... 91 C.11 Optimal Four and Five Stage Launch Vehicles according to the Pro- pellant Type∗ to Reach Semi-Synchronous Orbit...... 91 C.12 Optimal Launch Vehicles with Boosters Configuration according to the Propellant Type∗ to Reach Semi-Synchronous Orbit...... 91 C.13 Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Molnyia Orbit ...... 92 C.14 Optimal Three Stage Launch Vehicles according to the Propellant Type∗ to Reach Molnyia Orbit...... 92 C.15 Optimal Four and Five Stage Launch Vehicles according to the Pro- pellant Type∗ to Reach Molnyia Orbit...... 93 C.16 Optimal Launch Vehicles with Boosters Configuration according to the Propellant Type∗ to Reach Molnyia Orbit...... 93

ix

Nomenclature

∆V Real Delta-V

∆VD Drag Loss

∆VG Gravity Loss

∆Videal Ideal Delta-V

∆Vsteering Steering Loss

m˙ Propellant Mass Flow

γ Flight Path Angle

µ Lagrange Multiplier

µE Earth Gravitational Constant

ωE Earth Angular Velocity

ψ0 Ideal Azimuth

ψR Real Azimuth

CD Drag Coefficient

D Aerodynamic Drag

g0 Acceleration of Gravity at Sea Level

i Orbit Inclination

Isp Specific Impulse

xi xii NOMENCLATURE

kt Tank Contribution

L Latitude q Dynamic Pressure ra Apogee Radius rp Perigee Radius

RE Earth Radius

S Reference Area

T Thrust tb Flight Time

V0 Earth Rotation Velocity

Vf Orbit Velocity

Weng Engine Weight

xii Abstract

Since the Soviet Union launched the Sputnik (the first artificial satellite) in 1957, the Space Race became the major priority to both USSR and the US, making it possible to step the Moon in 1969 with the Apollo Missions. It was one of the most ambitious and expensive projects in history, where rockets had a lot of importance, particularly The Saturn V and the Soyuz.

Nowadays, rockets are still important for space exploration, from launching Cube- sats to LEO to solar system’s exploration. As the Space Race grew, new businesses related to rockets appeared. Within this thesis, it is studied the feasibility of a new business called "Rocket launching Consultancy" through the analysis of the propulsion system’s requirements for different missions. In order to know if there are different optimal rockets for the same mission.

xiii

Aim

The aim of this project is to study the requirements of rocket’s propulsion system for a few missions with different payloads up to 1,000 kg, to reach different LEO and MEO orbits, in order to determine how many launch vehicles can perform a specific mission and which one is the most sustainable rocket for that mission.

Finally, it is aimed at studying the feasibility to set up a business called "Rocket launching Consultancy", based on the propulsion system’s requirements for the different missions.

xv

Scope

Within this study, the topics listed below will be covered:

• Brief Market Research.

• State of the art of different payloads carried by launch vehicles, focusing on their weight.

• State of the art of actual ground launch sites.

• State of the art of operating launch vehicles.

• Definition of different missions based on the state of the art.

• Study of the rocket’s propulsion system requirements for each mission.

• Comparison between the obtained results and actual launch vehicles.

• Study of the feasibility of setting up a business

The following tasks are NOT within the scope of this study:

• Study of interplanetary orbits.

• Development of a Business Plan.

• Consideration of plane change manoeuvres.

xvii

Requirements

The requirements of this study are the following:

• The payload must not exceed 1,000 kg.

• At least three different LEO must be studied.

• At least one MEO must be studied.

• The studied payload must be divided into different weight ranges.

• It must be taken into account from which site each rocket can be launched.

• Only ground launch sites must be considered.

xix

Background

Nowadays, there is a wide variety of launch vehicles, each one with different fea- tures. It must be taken into consideration that their main function is to carry a payload to space, however, manufacturers only provide little data about the payload and the orbit that the rocket can reach. Therefore, due to this lack of information, customers might choose to launch their payload with a rocket that is oversized for that mission. Moreover, both launch costs and environmental impact would be higher, because it would be necessary more propellant.

According to [24], Small Satellite market is growing and it may surpass USD 62 Billion by 2030. This growth would entail small businesses start launching their satellites to space. However, because of their lack of resources, these small businesses would need someone to provide them advice about which launch vehicle would be the best option according to their necessities.

In order to get that market niche, which consists of advising the customers about the optimal launch vehicle, a new business idea called "Rocket launching Consul- tancy" was considered. Within this thesis it is tried to study the feasibility of this business, just considering a few different missions. It must be taken into consid- eration that this is an academic project, however, this study also could be used by some investors that would be interested in setting up a business similar to the proposed in this thesis. Moreover, this study also could be used to know if there are optimal launch vehicles that focus on these types of space missions, allowing the manufacturers to consider the design of a launch vehicle, specifically designed for one of those missions.

xxi

Chapter 1

Introduction

Since the Soviet Union launched the Sputnik (the first artificial satellite) in 1957, the Space Race became the major priority to both USSR and the US, making it possible to step The Moon in 1969 with the Apollo Missions. It was one of the most ambitious and expensive projects in history, where rockets had a lot of importance, particularly The Saturn V and the Soyuz.

Nowadays, rockets are still important for space exploration, from launching Cube- sats to LEO to solar system’s exploration, and every day more and more satellites are launched to outer space. As the Space Race grew, new businesses related to rockets appeared. Currently, it is forecast that by 2025 space launch services market would reach a value of USD 27.18 Billion according to [25].

Within this thesis, it will be studied the feasibility of new a business called "Rocket launching Consultancy” through the analysis of the propulsion system’s require- ments. But first of all, the services provided and how the feasibility study must be explained in more detail.

1 2 CHAPTER 1. INTRODUCTION

1.1 Brief Rockets’ Historical Review

Even though rockets were born in the 13th century when the Chinese and the Mongols were at war, rocketry did not become a science until the 17th century, when Isaac Newton was able to know more details about rockets, for example, their possibility to work in vacuum or outer space[26].

The idea of using rockets for space exploration was born in science-fiction books in the middle of the 19th century. Voyage to Venus, written by Achille Eyraud in 1965, was the first novel to describe an interplanetary rocket-powered spaceship[27]. Besides that book, there are other books that share the idea of using rockets to escape the earth. Examples of those are De la Terre à la Lune Trajet direct en 97 heures (Jules Gabriel Verne, 1965) or The First Men in the Moon (Herbert George Wells, 1902).

Nikolai Kibalchich, who was an explosive expert, in 1981 designed a kind of aero- nautic missile that would use a powder rocket engine. He was in prison when he thought about this concept, and he decided to send this design to the government, however, he did not get any response before being executed the same year.

It was not until the end of the 19th and the beginning of the 20th century when the idea of using rockets for space exploration became a reality due to Konstantin Tsiolkovsky. His idea consisted of using liquid propellants to achieve a greater range. On the other hand, early in the 20th century, Robert H. Goddard be- gan to perform practical experiments, firstly with solid fuel and later with liquid propellant, where he was able to achieve the first flight using liquid propellant. Another great pioneer was Hermann Oberth, due to him, rocket societies sprang up, including Verein fur Raumschiffahr, which one developed the V-2 and used it as a weapon during World War II[26].

After World War II and during the Cold War, both the United States and the Soviet Union began the Space Race[28] and started using rockets for space ex- ploration instead of using them as a weapon. Making possible to step the Moon in 1968 using the Saturn V, the most powerful rocket ever built[29]. Since then, rockets have been used for carrying spacecraft to outer space, and nowadays, they are still important for space exploration.

2 1.2. BUSINESS DESCRIPTION AND FEASIBILITY ANALYSIS ESEIAAT

1.2 Business Description and Feasibility Analysis

The business would consist of a kind of consultancy, which would indicate the customers the best launch vehicle and best launching site to launch their payload to space. In order to do that, the customer should provide some data about the payload (dimensions, mass, type, etc.) and also the orbit that they want to reach.

The function of this business would consist of selecting the optimal launch ve- hicle according to their conditions within a catalogue, which would include all active launch vehicles, not only a reduced fleet, as some launch services do. Later, knowing which one is the best option, the customer would be helped to book the launch, and also with regulatory paperwork. Unlike companies that provide launch services, this business would not provide support about technical phases of the mission.

This would be a digital consultancy, where all the procedures would be done through Internet. This would allow customers from all around the world to contact us with no complication through e-mail, the consultancy’s web page, or video call.

Regarding the feasibility study, the main concern is to know if there are a lot of optimal launch vehicles for a specific mission. It is considered that the cus- tomer would need advice if there are not too many launch vehicles that satisfy the requirements because a deeper study should be done in order to find optimal rockets. On the other hand, if there is a wide variety of optimal launch vehicles that can perform the same mission, with a more basic study would be enough to select a launch vehicle that adapts to the mission. In order to get an approach to the feasibility of the business, a few different missions with different features mainly focused on small satellites (because of their advantages and the growth of small satellites market) are studied to verify if there are a lot of options for each mission.

To determine which would be the optimal launch vehicles, the propulsion system’s requirements for different missions are calculated, because just analysing the data given by manufacturers, a wide variety of missions would be encompassed, and it would be more difficult to compare the different options.

3 4 CHAPTER 1. INTRODUCTION

1.2.1 Potential Customer

This consultancy would provide a service that would be oriented towards private companies that need to launch satellites, mainly small satellites. These companies should not have enough resources to make a study about launch vehicles, neither having a previous agreement with some launch services provider.

As said, the small satellite market will grow in the following years, creating new businesses that would require launch services. Therefore, this consultancy would intend to satisfy the necessities of those new companies that would not have a large budget at the beginning, and they would need an easy way to save money throughout the different phases of their projects.

These would be the potential customers during the first years. Although, as time goes by, it could be interesting to get agreements with bigger companies, which would allow obtaining more benefits, however, this idea is too ambitious at the beginning.

4 Chapter 2

Market Research

Taking into account that this project includes a business idea, studying its market niche is mandatory. Within this chapter, it is explained both, how customers decide which launch service to buy and also potential competitors.

2.1 Buying a Launch Service

The whole process from buying a launch service to the final launch can take months or even years [30]. In order to get the payload into the orbit, it is necessary to get throughout different phases, one of them is buying a launch service. This phase can be harder than expected because the launch vehicle and the carried spacecraft must be compatible. However, there a lot of factors that must be taken into account, for example, launch vehicle and spacecraft’s features or destination requirements [11].

As it has been said, buying a launch service can be difficult, nevertheless, two different paths can be taken. The first one would consist of contacting directly to different manufacturers. This would take more time and personal implications, however, it would be probably cheaper. On the other hand, the second option would consist of using something, usually a company, that acted as a broker and matched the spacecraft with the optimal launch vehicle [11]. Unlike the first path, this one would involve a greater cost just by the fact of the existence of an

5 6 CHAPTER 2. MARKET RESEARCH intermediary. To get more accurate results about the cost of each path, it would be necessary to do a deeper study, which is beyond the scope of this project.

2.2 Competitors

First of all, it is necessary to know which would be our place in the market. As it was explained in Sec. 1.2, this business idea would consist of providing advice to customers, therefore, our function would be to act as a broker. Currently, there are other companies that are already providing this kind of service. They are listed below:

2.2.1 NASA Launch Services Program

Established in 1998 at Kennedy Space Center, NASA’s Launch Services Program was created to manage support to spacecraft customers until the present day. LSP is responsible for matching spacecraft with optimal launch vehicles, selecting the right one. Then, LSP buys that spacecraft a ride to space in order to ensure mission success. Along this process, from pre-mission planning to spacecraft’s post-launch phase, LSP provides support. All this, always keeping in mind safety, reliability, cost-effectiveness, and on-schedule processing. Besides this full service, LSP also provides tailored approaches and advisory services to different customers[11].

This whole service consists of different phases, which are listed below [11]:

• Advanced Planning: within this phase it is provided support to spacecraft design and trade study of launch vehicles.

• Business: mission’s budget is managed within this phase.

• Technical: within this phases verification, validation, certification and space- craft integration are provided in order to achieve mission’s goals.

• Launch Site Operations: within this phase, infrastructure to communi- cate with spacecraft id provided.

• Launch Operations: communication and telemetry data is provided. 6 2.2. COMPETITORS ESEIAAT

• Post launch: mission success is determined within this last phase.

In order to select the optimal launch vehicle for spacecraft’s requirement, LSP offers a fleet of different launch vehicles, which can be seen in Table 2.1.

Vehicle Class Small Medium Intermediate/Heavy Falcon 9 Full Launch Vehicle Pegasus XL Taurus-XL Antares 2XX Atlas V 4XX Atlas V 5XX Delta IV Heavy Falcon Heavy Thrust Offeror OSC OSC OSC SpaceX ULS ULS ULS SpaceX CCAFS CCAFS WFF CCAFS CCAFS CCAFS CCAFS Launch Sites WFF WFF CCAFS KWAJ VAFB VAFB VAFB VAFB VAFB VAFB

Table 2.1: LSP Fleet for Launching Payloads into LEO or beyond [11]

2.2.1.1 NASA Venture Class Launch Services

This was fund by LSP and it is designed for Cubesat missions, providing faster launch services[31]. It is used to launch small satellites, therefore, smaller launch vehicles are used. Nowadays, only Electron (Rocket Lab) and LauncherOne (Vir- gin Orbit) are under the contractual mechanism of VCLS [11].

2.2.2 Spaceflight

Founded in 2011, Spaceflight provides end-to-end launch services and also provides rideshare services to get lower launch cost[1].

Its services are divided into different phases[32]:

• Planning: within this phase the optimal launch vehicle is determined ac- cording to customer’s budget and timeframe.

• License: all licensing requirements for launch are addressed within this phase.

• Integrate: all the process to integrate the payload is addressed within this phase in order achieve mission’s goals.

• Transport: advise is provided to customers about the best way to get the payload to the launch site. 7 8 CHAPTER 2. MARKET RESEARCH

• Launch: the customer will be guided through the whole launch day.

In order to select the optimal launch vehicle, Spaceflight is manifesting missions with a wide variety of launch vehicles, which can be seen in Fig. 2.1:

Figure 2.1: Spaceflight Launch Vehicle Family[1]

2.2.3 Precious Payload

Founded in 2017, Precious Payload is a digital service for satellite launch plan- ning. It offers services related to management and control over mission progress, assistance to evaluate mission cost and management with all space regulatory spacework. It also provides help through all the steps to get the payload into orbit, which includes the following topics [33]:

• Finding launch options

• Creating mission roadmap

• Finding ground segment options

• Regulatory compliance

• Buying insurance 8 2.2. COMPETITORS ESEIAAT

Precious Payload handles rideshares slots posted by launch operators and other services from other suppliers, all this, aiming at reducing costs and time associated with mission management by the implementation of matchmaking algorithms [34].

2.2.4 Loft Orbital

Founded in 2017 in San Francisco, Loft Orbit offers an end-to-end service that delivers the mission to orbit on a standard microsatellite bus. Its service includes satellite design, licensing and regulation, booking a launch, ground segment, and financing and insurance. Loft Orbit aims at making space simple[35].

2.2.5 EXOLAUNCH

Founded in 2008, EXOLAUNCH is a European launch services, mission man- agement, and deployment system provider for small satellites. It provides an end-to-end service that includes launch planning, project management, separation systems, licensing, testing, shipment, integration, and launch[36].

2.2.6 D-Orbit

Founded in 2011 in Como (Italy), D-Orbit intends to cover the entire life-cycle of a space mission, including mission analysis and design, engineering, manufacturing, integration, testing, launch, and end-of-life decommissioning. It focuses on small satellite missions and deployment service[37].

2.2.7 Other Launch Services Providers

Besides to these companies, there are other launch services providers that offer end-to-end launch services. However, these have smaller fleets of launch vehicles, making more difficult the selection of an optimal rocket. Some of these launch services providers are listed below [38]:

• Antrix Corporation 9 10 CHAPTER 2. MARKET RESEARCH

• Arianespace

• China Aerospace Science and Technology Corporation

• International Launch Services

• Indian Space Research Organisation

• Orbital ATK

• SpaceX

• United Launch Alliance

Analysing all competitors mentioned in the sections above, it can be seen that almost everyone provides end-to-end services, which takes care of every aspect of the mission (from planning to post-launch). However, the business proposed in this project has a different view, which focuses on a specific part of the mission, the planning.

This business intends to solve one of the main problems of launching a payload into space, which is finding a launch vehicle that satisfies its necessities. This service would be cheaper in comparison to the competitors presented in this chapter, making it more affordable for small businesses.

While current competitors have contracts with a few launch services providers, this consultancy would intend to have as many contracts as possible in order to select the best launcher according to customers’ necessities. Once the launcher is selected, the customer would be lead to contact with the launch service provider. The consultancy also would assist in regulatory paperwork that could be necessary to launch the payload.

In conclusion, this business idea intends to make satellite launching more affordable through customer guidance throughout the planning phase.

10 Chapter 3

State of the art

Within this chapter, all parameters, as the payload, orbits, launch sites and op- erating launch vehicles, needed for the further study of the propulsion system’s requirements are explained.

3.1 Payload

Launch vehicles are used to carry payload from Earth’s surface to Earth’s orbits or beyond. There are different types of payload, which are defined in the following section.

3.1.1 Human Spacecraft

These spacecraft are used for space travelling with crew aboard. They must meet countless requirements [39] in order to guarantee crew safety and mission achieve- ment.

Currently, there are only two operating human spacecraft, the Soyuz and the Shenzhou. On the one hand, the Soyuz is a Russian spacecraft that carries people and supplies to the ISS, and it can also bring them back [40]. On the other hand, the Shenzhou is a Chinese spacecraft, however, it is based on the Soyuz [41]. There

11 12 CHAPTER 3. STATE OF THE ART are other spacecraft under development, which are listed in Table 3.1 alongside their specifications:

Type Soyuz[42] Shenzhou[41] Dragon 2[43] CST-100[44] Orion[45]

Mass [kg] 7,150 7,840 9,525 13,000 35,385 Crew 3 3 7 7 4 Capsule Volume [m3] 8.5 14.0 9.3 11.0 8.95

Table 3.1: Human Spacecraft Specifications

As can be seen, all these spacecraft exceed 1,000 kg, which is the maximum payload that can be studied within this project. Moreover, each one of these spacecraft only can be launched with a single or few models of launch vehicles, for example, the Soyuz capsule only can be launched with Soyuz rocket. Thus, this type of payload is not used for the study carried out in Chapter 4.

3.1.2 Cargo Spacecraft

This type of payload is similar to human spacecraft explained in Sec. 3.1.1 except by the fact that they are designed to carry cargo instead of carrying people to space. They are mainly used for resupplying the International Space Station and also to bring cargo back to Earth [46].

In the same way, as in the section above (Sec. 3.1.1), there are only a few spacecraft that can perform these missions. They are presented in Table 3.2:

Type Progress HTV Dragon[47] Cygnus Tianzhou[48]

Dry mass [kg] 5,740 11,000 4,200 1,500 7,000 Payload [kg] 1,700 5,500 6,000 1,700 6,500 Return payload to Earth [kg] 0 0 3,000 0 -

Table 3.2: Cargo Spacecraft Specifications [12]

Analysing this type of payload, it is clear that it exceeds 1,000 kg, which is the maximum allowed according to the requirements of this thesis. Thus, as happened with human spacecraft, this type of payload is not used for further studies.

12 3.1. PAYLOAD ESEIAAT

3.1.3 Space Probes

According to [49], a space probe is an unmanned spacecraft. It travels through space, and it is also able to orbit or land on a planet or a moon to collect science information. The acquired data is sent back to Earth for scientific studies.

Since Sputnik 1 was launched in 1957, a wide variety of probes have been sent to space. Some of them are listed in Table 3.3:

Type Voyager 1 Cassini-Huygens Juno Insight Mars Odyssey

Mass [kg] 815 5,712 3,625 694 725 Destination Out of solar system 5,500 Jupiter Mars Mars Launched 1977 1997 2011 2018 2001

Table 3.3: Space Probes Features [13][14]

This type of spacecraft can weigh less than 1,000 kg, complying with the require- ments of this thesis regarding the maximum payload. However, they are designed to travel through space, away from Earth’s orbits. Even though these spacecraft could orbit Earth, this type of payload neither is considered for further studies within this thesis.

3.1.4 Satellites

According to [50], a satellite is an object that moves around a larger object. Those objects can be natural satellites, as for example the moon, or artificial satellites, which are man-made. In this section, only artificial satellites that orbit Earth will be taken into consideration. Satellites can be classified according to their size or their applications.

Different satellites applications are listed below [51]:

• Navigation: these satellites are used to provide the user 24-hour three- dimensional positions, velocity and time information on Earth’s surface.

• Communication: these satellites are used for telecommunication applica- tions, such as telephony, television or radio. 13 14 CHAPTER 3. STATE OF THE ART

• Weather: these satellites are used for different meteorological observations.

• Earth observation: these satellites are used to understand and to analyse environmental conditions.

Satellite classification according their mass is presented in Table 3.4:

Type Mass

Large Satellites >1,000 kg Medium Satellites 500 to 1,000 kg Minisatellites 100 to 500 kg Microsatellites 10 to 100 kg Nanosatellites 1 to 10 kg Picosatellites 0.1 to 1 kg Fenitosatellites <0.1 kg

Table 3.4: Satellite Classification according to Mass [15]

Satellites above do not have a specific geometry. However, since 1999 there is a class of nanosatellites called CubeSats, which use a standard size called "one unit (1U)". This unit measures 10x10x10 cm and weighs 1.33 kg. Cubesats can extend to larger sizes, such as 1.5, 2, 3, 6 or even 12U[52].

3.2 Rocket Launch Sites

Nowadays, rockets can be launched from different platforms, including ground- based platforms, air-launched vehicles, mobile sea-based platforms or even from submarines. However, within this section, only ground-based platforms are con- sidered, in order to comply with the requirements.

14 3.2. ROCKET LAUNCH SITES ESEIAAT

3.2.1 Geographic Considerations

In order to place a satellite into orbit, high altitude and high horizontal velocity are required. Due to Earth’s shape and its rotation around its centre axis, some positions are more optimal than others[2].

3.2.1.1 Latitude

Earth’s natural horizontal surface velocity depends on latitude, having higher ve- locities at lower latitudes. Furthermore, this velocity is eastwards because Earth rotates west to east. Thus, to take advantage of Earth’s rotation, the launch must be performed due to eastward from the equator. However, not all missions require high velocities, therefore, for those, lower latitudes are not the most optimal.

Figure 3.1: Velocity at Earth’s Surface by Latitude [2]

3.2.1.2 Azimuth

According to [2], rocket’s launch azimuth is the direction it travels in the horizontal plane after leaving the launch pad, measured in degrees clockwise from due north. All launch sites have drop zones, where rocket stages fall back to Earth during a successful mission or aborted mission. In case of a crash, different elements of a rocket could be hazardous for people. This is the reason why launch sites have 15 16 CHAPTER 3. STATE OF THE ART azimuth limitations, in order to avoid creating drop zones that include populated areas or foreign territory.

Within this section, the parameters that must be taken into consideration for rocket launch sites have been explained, whereas, in the following section (Sec. 3.2.2), all active spaceports and their features are presented.

3.2.2 Active Launch Sites

Nowadays, there are 22 active spaceports around the world and all of them are listed in Table 3.5:

Launch site Latitude Longitude Max Azimuth Min Azimuth Other Considerations Plesetsk Cosmodrome 62.9oN 40.6oE 90o -19o - Baiknour Cosmodrome 46.0oN 63.3oE 65o -13o - Cape Canaveral/KSC 28.6oN 80.6oW 120o 35o - Vandenberg Air Force Base 34.6oN 120.6oW 201o 158o - Guiana Space Centre 5.2oN 52.8oW 92.5o -10o - Xichang Satellite Launch Center 28.3oN 102.0oE 104o 94o - Jiuquan Satellite Launch Center 41.0oN 100.3oE 153o 134o - Tanegashima Space Center 30.4oN 131.0oE 180o 0o - Taiyuan Satellite Launch Center 38.9oN 111.6oE 192o 175o - Satish Dhawan Space Centre 13.7oN 80.2oE 140o 0o - Wallops Flight Facility 37.9oN 75.5oW 160o 90o - Uchinoura Space Center 31.3oN 131.1oE 155o 25o - Yasny Launch Base 51.1oN 59.8oE 45o -13o - It is classified, however, this spaceport Westward Westward Palmachim Airbase 31.9oN 34.7oE only performs westward launch in order (see Fig. 3.3) (see Fig. 3.3) to avoid adversaries to east[2]. This spaceport is operated by the Iranian Space Agency. All successful Imam Khomeini Space Center 35.2oN 54.0oE - - orbital launches reached LEO at inclinations between 55o and 56o[2]. Vostochny Cosmodrome 51.9oN 128.3oE 90o -13o - From this site it is possible to reach orbital Rocket Lab Launch Complex 39.3oS 177.9oE approx. 350o approx. 90o inclinations from sun-synchronous through to 39 degrees[53]. Pacific Spaceport Complex - Alaska 57.4oN 152.3oW 220o 110o - Southward Southward This spaceport is operated Sohae Satellite Launching Station 39.7oN 124.7oE (see Fig. 3.3) (see Fig. 3.3) by North Korea[2]. Wenchang Satellite Launch Center 19.6oN 111.0oE 190o 90o - This base is used for missile testing, Ronald Reagan Ballistic Missile however, it also has hosted commercial 9.1oN 167.7oE See Fig. 3.2 See Fig. 3.2 Defense Test Site launches for companies like SpaceX and Orbital Science[3]. Naro Space Center 34.4oN 127.5oE 190o 90o -

Table 3.5: Active Orbital Launch Sites [2][16]

16 3.3. OPERATING LAUNCH VEHICLES ESEIAAT

It must be taken into account that these sites are for orbital launches, whereas sub-orbital sites are not considered. All this information is complemented with the data from Sec. 3.3, where it is deter- mined from what site each rocket can be launched. Both are used for the study carried out in Chapter 4.

Figure 3.2: Potential Launch Az- imuths from Omelek[3]

Analysing Table 3.5, there are three launch sites, whose maximum and minimum azimuths are not indicated because there is no information about them. In order to get an approximation about values of Palmachin and Sohae, the following diagrams have been done:

Figure 3.3: Palmachim and Sohae Hypothetical Azimuths

3.3 Operating Launch Vehicles

Currently, the only way to carry a payload to Earth’s orbits or beyond is using rockets. Within this section, a brief explanation about rocket features is done. In addition, all current operating launch vehicles along their features are also presented. 17 18 CHAPTER 3. STATE OF THE ART

3.3.1 Rocket Features Definition

Rockets have a wide variety of characteristics, however, within this section, only a few of them are defined, which are the most relevant for the development of this thesis.

Stages

Most of the mass of a rocket is due to propellant mass, whereas the payload is a small portion of lift-off mass. In order to achieve orbital velocity, most launch vehicles discard a portion of the vehicle when it is run out of propellant, lightening the weight of the rocket[4]. This process is called staging, therefore, stages could be defined as the different portions of a launcher that separates when the propellant is run out. Rockets can have a different number of stages, and they also can use boosters (Fig. 3.5) to generate more thrust.

Figure 3.4: Serial Staging [4] Figure 3.5: Parallel Staging [4]

Thrust

The thrust is the force generated by a rocket propulsion system acting upon a vehicle. This force is due to the ejection of matter at high velocity and pushes the rocket through the application of Newton’s third law of motion[4].

18 3.3. OPERATING LAUNCH VEHICLES ESEIAAT

Specific impulse

The specific impulse is an important figure of merit of the performance of a rocket propulsion system and also it is used to compare propellants or engine efficiency. Likewise thrust, it is higher in vacuum than at sea level[54].

Propellant

Regarding propellants, it is necessary to differentiate between the fuel and the oxidiser and also between solid and liquid propellants. Now, let’s focus only on the fuel and the oxidiser. In order to get a first approach, the fuel is a substance that burns in the presence of oxygen, whereas the oxidiser is a source of oxygen[55]. Propellants are studied in more detail in Sec. 4.3.3.

3.3.2 Launch vehicles

Knowing what rocket’s parameters mean, it is time to look for operating launch vehicles, which are necessary in order to make comparisons later in Chapter 5.

Currently, there are nearly 90 different orbital launchers operating around the world, including all variants of a family[21]. As this is a large list, they are pre- sented along with some of their specifications in Annex A, whereas launch vehicles and their launch sites are presented in Table 3.6. It must be outlined that are listed only the sites that have at least one launch vehicle operating. Thus, nei- ther Ronald Reagan Ballistic Missile Defense Test Site, nor Naro Space Center are included.

19 20 CHAPTER 3. STATE OF THE ART

Two Stages Three Stages Four Stages Five Stages Launch site Launch Vehicles Launch Vehicles Launch Vehicles Launch Vehicles Rochot Soyuz 2.1a/b Plesetsk Cosmodrome Angara - - Soyuz 2.1v Zenit Dnepr Proton Medium Baiknour Cosmodrome - Proton M Soyuz 2.1v Strela Alpha Atlas V Cape Canaveral Air Force Station Beta Minotaur IV - Minotaur-C Delta IV Minotaur V Kennedy Space Center Falcon 9 Falcon Heavy Alpha Atlas V Beta Minotaur IV Delta IV Minotaur I Vandenberg Air Force Base - Minotaur V Falcon 9 Minotaur-C Minotaur VI Falcon Heavy Vector-H Vector-R Vega Guiana Space Centre Ariane 5 Soyuz 2.1a/b - Vega-C Long March 2C Long March 3A Xichang Satellite Launch Center - - - Long March3B/E Long March 3C Long March 2C Hyperbola-1 Long March 2D Long March 4B Jiuquan Satellite Launch Center Kuaizhou 1 - Long March 2F Long March 4C Kuaizhou 11 Zhuque-1 Tanegashima Space Center H-IIA/B - - - Long March 2C Long March 4B Taiyaun Satellite Launch Center - - - Long March 4C Long March 6 PSLV Satish Dhawan Space Centre LVM3 GLSV - SSLV Minotaur IV Vector-H Minotaur I Wallops Flight Facility (MARS) Antares Minotaur V Vector-R Minotaur-C Minotaur VI Uchinoura Space Center - - Epsilon - Yasny Launch Base - Dnepr - - Palmachim Airbase Shavit 2 - - Imam Khomeini Space Center - Safir - - Vostochny Cosmodrome Angara - - - Rocket Lab Launch Complex Electron - - - Minotaur IV Vector-H Pacific Spaceport Complex - Alaska - Minotaur-C Minotaur V Vector-R Minotaur VI Sohae Satellite Launching Station - Unha - - Long March 5 Wenchang Satellite Launch Center - - - Long March 7

Table 3.6: Launch vehicle’s sites 20 3.4. ORBITS ESEIAAT

3.4 Orbits

According to [56], an orbit is a regular, repeating path that one object in space takes around another one. Within this thesis, it is focused on Earth’s orbits, which can be divided into different types. But first of all, it is necessary to introduce the parameters that define an orbit.

3.4.1 Orbit Parameters

Orbits are elliptical in shape, therefore, the object, in this case, satellites are not always the same distance from Earth.

Figure 3.6: Orbital Elements [5]

The elements used to describe the motion of a satellite within an orbit are repre- sented in Fig. 3.6 and summarised below[5]:

• Semi-major axis (a): defines the size of the orbit.

• Eccentricity (e): defines the shape of the orbit.

• Inclination (i): defines orbit’s orientation respect to the Earth’s equator.

• Argument of Perigee (ω): defines where the low point, perigee, of the orbit is with respect to the Earth’s surface.

21 22 CHAPTER 3. STATE OF THE ART

• Right Ascension of the Ascending Node (Ω): defines the location of the ascending and descending orbit locations with respect to the Earth’s equatorial plane.

• True anomaly (ν): defines where the satellite is within the orbit with respect to perigee.

As said, these parameters are needed to define the motion of the satellite. However, in order to define the target orbits, which are required for the development of this thesis, it is enough defining the semi-major axis, the eccentricity and the inclination.

3.4.2 Earth orbit classification

Orbits can be classified into different groups according to their characteristics. In this section, orbits are classified according to their altitude. This classification is presented below[57]:

Low Earth Orbit (LEO)

These orbits have a range of altitudes from 160 km up to 2,000 km above Earth’s surface. Due to their proximity to Earth, these orbits are used for imaging, remote sensing or military purposes[58].

Medium Earth Orbit (MEO)

These orbits take place at an altitude of 2,000 km, up to 35,788 km approximately, where it is found the geosynchronous orbit. These orbits are normally used for communications or navigation satellites.

Geosynchronous Earth Orbit (GEO)

This orbit has been defined to be within a 600 km range, from 35,488 km up to 35,788 km. Satellites located in this orbit are used for communications because 22 3.4. ORBITS ESEIAAT of the fact that satellites seem to be at the same location because they have the same orbital period as the Earth[58].

High Earth Orbit (HEO)

This orbit is above the geosynchronous orbit. Satellites in this orbit have an or- bital period longer than the one of the Earth, which is 23 h 56 min 4.09 s.

Figure 3.7: Earth Orbits Representation

23

Chapter 4

Study of Propulsion System’s Requirements

In this chapter, all calculations required to carry out this study are done. Starting from defining all missions that are considered within this study, to the calculation of the propulsion system’s parameters for each one of the missions.

4.1 Missions Parameters and Definition

In order to define the missions and also complying with the requirements of this thesis, it is necessary to define the different ranges of mass that are considered and the different orbits to reach.

Firstly, it is started with the payload. Taking into consideration that within this study the payload only is defined by its mass, the classification presented in Table 3.4 is used to define the studied payload. Thus, the different ranges of mass are the following:

• Payload Type I : 1,000 kg.

• Payload Type II : 500 kg.

• Payload Type III : 100 kg.

25 26 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

• Payload Type IV : 10 kg.

Finally, it is necessary to define the target orbits for this study. According to the requirements, it is necessary to define at least three LEO and one MEO. Depending on orbits’ features, they can be considered direct ascent missions or parking orbit ascent missions.

According to [9], direct ascents are used for Low-Earth circular orbits or elliptic orbits easily reached from the launch site. Whereas parking orbit ascent is mainly used for high LEO, MEO and Geosynchronous Transfer Orbit, which require or- bital manoeuvres to reach them.

Most of the missions get through a parking orbit, which is located 200 km above sea level approximately (see Fig. 4.1 ). This parking orbit is taken as reference for the orbital manoeuvres performed to carry out the development of this project. Below are shown the final target orbits along with their parking orbits, that are considered in this thesis.

(a) Soyuz [59] (b) H-II [19]

Figure 4.1: Launch Vehicle Performance and Launch Mission.

Low Earth Orbits

Below are listed three Low Earth Orbits, with different characteristics from each other:

Polar orbit

This type of orbit passes over Earth’s polar regions. It is also called polar orbit, even though it passes within 20 to 30 degrees of the poles. One example could be the orbit that Globalstar satellites follow: 26 4.1. MISSIONS PARAMETERS AND DEFINITION ESEIAAT

• Perigee (rp) = 1,500 km

• Apogee (ra) = 1,500 km

• Inclination (i) = 82.5o

Considering this as high LEO, the spacecraft is placed into a transfer orbit, whose parameters are listed below, in order to finally reach the target orbit.

• Perigee Parking Orbit (rp) = 200 km

• Apogee Parking Orbit (ra) = 1,500 km

• Inclination Parking Orbit (i) = 82.5o

Sun-Synchronous Orbit

The main advantage of this orbit is the possibility to observe the same point with the same lighting conditions. One example of this orbit could have the following features:

• Perigee (rp) = 705 km

• Apogee (ra) = 705 km

• Inclination (i) = 98.2o

This orbit also could be considered as a high LEO, therefore, a parking orbit is used to reach it:

• Perigee Parking Orbit (rp) = 200 km

• Apogee Parking Orbit (ra) = 705 km

• Inclination Parking Orbit (i) = 98.2o

Inclined Orbit

A clear example of this orbit could be the International Space Station (ISS), whose orbit has the following parameters: 27 28 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

• Perigee (rp) = 415 km

• Apogee (ra) = 415 km

• Inclination (i) = 51.2o

The parking orbit to reach this orbit would have the following parameters:

• Perigee Parking Orbit (rp) = 200 km

• Apogee Parking Orbit (ra) = 415 km

• Inclination Parking Orbit (i) = 51.2o

It must be taken into consideration that in Low Earth Orbits, spacecraft experience atmospheric drag, making the spacecraft descend over time. Moreover, Earth’s shape is not regular. Therefore, the distance between satellites and the Earth can vary along their lifetime. However, this aspect is not necessary to take into account for the development of this thesis, considering that the satellite would be placed in the desired altitude when it is launched to space, and orbital maintenance is out of the scope of this thesis.

Medium Earth Orbits

There are not a lot of satellites in Medium Earth Orbit. Thus, it is intended to define an orbit similar to current MEO satellites already have.

Semi-Synchronous Orbit

This type of orbit has the particularity of having an orbital period of approximately 12 hours. In this orbit, it can be found navigation satellites such as Global Posi- tioning System (GPS), GLONASS or Galileo. In order to study it, the considered target orbit has the following features:

• Perigee (rp) = 21,000 km

28 4.1. MISSIONS PARAMETERS AND DEFINITION ESEIAAT

• Apogee (ra) = 21,000 km

• Inclination (i) = 60.0o

Navigation Apogee [km] Perigee [km] Inclination [o] Satellite

GLONASS 19,140 19,140 64.8 GPS 20,180 20,180 55 GALILEO 23,222 23,222 56

Table 4.1: Difference between GLONASS, GPS and Galileo [17]

In order to reach this orbit, orbital manoeuvres are required. The spacecraft is placed into a transfer orbit, whose parameters are the following:

• Perigee Parking Orbit (rp) = 200 km

• Apogee Parking Orbit (ra) = 21,000 km

• Inclination Parking Orbit (i) = 60.0o

Through this orbit, it would be possible to reach the desired target orbit.

Molniya Orbit

This an example of a highly elliptical orbit, which is used for communication satellites. Its parameters are the following:

• Perigee (rp) = 540 km

• Apogee (ra) = 39,300 km

• Inclination (i) = 63.4o

In this case, the parking orbit has the following parameters:

• Perigee Parking Orbit (rp) = 200 km

29 30 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

• Apogee Parking Orbit (ra) = 540 km

• Inclination Parking Orbit (i) = 60.0o

In order to cover a wide variety of missions within the limitations of this thesis, it is considered that the four different types of payload (I to IV) must reach each one of the orbits presented above. All this, making possible to have a total of 20 different missions.

4.2 Study of the Possibility to Reach the Orbits

In order to ease all calculations, this section is dedicated to studying if it is possible to reach target orbits from the different launch sites presented in Sec. 3.2.2.

It must be taken into consideration that rockets must be launched with a certain azimuth in order to reach a specific orbit. It depends on orbit’s inclination and the launch site according to Eq. 4.1:

cos i ψ = arcsin (4.1) 0 cos L

This angle does not take into account Earth’s Rotation, which can be defined as:

V0 = ωERE cos L (4.2)

Topocentric horizon Figure 4.2: Figure 4.3: Launch Azimuth[7] coordinate system [6]. A is the az- imuth.

30 4.2. STUDY OF THE POSSIBILITY TO REACH THE ORBITS ESEIAAT

To obtain the real azimuth it is necessary to obtain the orbit velocity, which is at

the perigee (r = rP ):

s 1 1 Vf = Vp = 2µE( − ) (4.3) r rp + ra

Having calculated all these parameters, it is possible to determine the ideal Delta- V necessary to reach the orbit (Eq. 4.4):

q 2 2 ∆Videal = Vf + V0 − 2Vf V0 sin ψ0 (4.4)

Finally, the real azimuth necessary to reach the orbit it is obtained from Eq. 4.5:

2 2 2 Vf − V0 − ∆Videal ψR = arcsin (4.5) 2V0∆Videal

As said in Sec. 4.1, for some missions it is necessary to perform orbital manoeu- vres to reach the target orbit. The most energy efficient two-impulse manoeuvre for transferring between two co-planar circular orbits sharing a common focus is the Hohmann transfer, which is explained by Wertz [Space Mission Analysis and Design, p.147 ].

Figure 4.4: Hohmann Transfer[6]

In this project, it is considered that rockets are launched directly to the elliptical orbit for those missions that require orbital manoeuvres. The necessary velocity

31 32 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

to go from one orbit to another is given by:

∆Vmanoeuvre = Vneed − Vcurrent (4.6)

All these steps have been followed for each one of the target orbits, and results are presented in sections below.

4.2.1 Possibility to Reach the Polar Orbit

As said, there are 22 active launch sites. However, in the table below are repre- sented those from where the target orbit can be reached, whereas in Annex B it can be seen the calculations for all sites.

o Launch Site ∆Videal [m/s] ψR [ ] Plesetsk 8,068.67 15.21 Baiknour 8,072.35 8.58 VAFB 8,074.96 173.56 Guiana SC 8,079.17 4.27 Tanegashima 8,075.85 5.89 / 174.11 Satish Shawan 8,078.53 4.54 Yasny 8,071.17 9.97 Vostochny 8,070.99 10.22 Mahia 8,073.91 172.81 PSCA 8,069.78 167.71 Sohae 8,073.82 172.74 Wenchang 8,077.78 175.12

Table 4.2: Delta-V to Reach the Parking Orbit (Polar Orbit)

From the elliptical orbit (perigee = 200 km and apogee = 1800 km), it is necessary an extra Delta-V to reach the circular orbit:

∆Vman = 327.4 m/s

32 4.2. STUDY OF THE POSSIBILITY TO REACH THE ORBITS ESEIAAT

4.2.2 Possibility to Reach the Sun-Synchronous Orbit

o Launch Site ∆Videal [m/s] ψR [ ] VAFB 8,002.1 192.68 Mahia 8,001.1 193.16 PSCA 7,996.9 197.08

Table 4.3: Delta-V to Reach the Parking Orbit (Sun-Synchronous Orbit)

From the elliptical orbit (perigee = 200 km and apogee = 705 km), it is necessary an extra Delta-V to reach the circular orbit:

∆Vman = 140.0 m/s

4.2.3 Possibility to reach the Inclined Orbit

o Launch Site ∆Videal [m/s] ψR [ ] Baiknour 7,556.5 63.37 CCAFS/KSC 7,560.7 43.37 Guiana SC 7,563.8 36.26 Tanegashima 7,560.3 44.50 / 135.50 Satish Shawan 7,563.1 37.55 MARS 7,558.5 129.12 Uchinoura 7,560.1 45.12 Imam 7,559.2 48.22 Mahia 7,558.2 127.53 Wenchang 7,562.3 140.79

Table 4.4: Delta-V to Reach the Parking Orbit (Inclined Orbit)

From the elliptical orbit (perigee = 200 km and apogee = 415 km), it is necessary an extra Delta-V to reach the circular orbit:

∆Vman = 61.8 m/s

33 34 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

4.2.4 Possibility to Reach the Semi-Synchronous Orbit

o Launch Site ∆Videal [m/s] ψR [ ] Satish Shawan 9,673.8 28.68 MARS 9,670.2 142.36 Uchinoura 9,671.4 33.91 / 146.09 Imam 9,670.7 35.94 Vostochny 9,667.5 53.13 Mahia 9,669.9 141.38 PSCA 9,666.4 112.42 Wenchang 9,673.1 150.14

Table 4.5: Delta-V to Reach the Parking Orbit (Semi-Synchronous Orbit)

From the elliptical orbit (perigee = 200 km and apogee = 21,000 km), it is neces- sary an extra Delta-V to reach the circular orbit:

∆Vman = 1, 442.8 m/s

4.2.5 Possibility to Reach the Molniya Orbit

o Launch Site ∆Videal [m/s] ψR [ ] Plesetsk 7,673.6 79.10 Baiknour 7,677.5 38.29 Guiana SC 7,684.7 23.63 Jiuquan 7,678.7 145.72 Tanegashima 7,681.2 28.72 / 151.28 Satish Shawan 7,684.0 24.45 MARS 7,679.5 147.68 Uchinoura 7,681.0 150.92 Yasny 7,676.3 43.95 Imam 7,680.1 30.85 Vostochny 7,676.1 45.05 Mahia 7,679.1 146.84 PSCA 7,674.8 124.83 Wenchang 7,683.2 154.50

Table 4.6: Delta-V to Reach the Parking Orbit (Molniya Orbit)

34 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

From the elliptical orbit (perigee = 200 km and apogee = 540 km), it is necessary an extra Delta-V to reach the final target orbit:

∆Vman = 2, 509.5 m/s

4.3 Propulsion System’s Parameters Calculation

Besides the target orbit, other aspects must be defined before obtaining the propul- sion system’s parameters needed for the different missions.

4.3.1 Delta V Budget

To describe the required energy to perform a space mission, it is used the mission velocity, also called ∆V . In a hypothetical case (gravity-free vacuum), the ∆V is calculated using Eq. 4.4, which takes into consideration the target orbit and Earth’s rotation. However, the propulsion system has to provide enough energy to also overcome the velocity losses, which appear due to the gravity effect, drag from the atmosphere and the necessity to steer the launch vehicle. Also considering the manoeuvres, the required mission velocity can be defined as:

∆V = ∆Videal + ∆Vman + ∆VD + ∆VG + ∆Vsteering (4.7)

Both, ∆Videal and ∆Vman have already been defined in Sec. 4.2, whereas the drag loss and the gravity loss are given by:

Z tf D Z tf ∆VD = dt ∆VG = g0 sin γ dt (4.8) t0 m t0

35 36 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

Figure 4.5: Launch Vehicle Boost Trajectory[8]

According to [8], these integrals cannot be computed because the drag, acceleration of gravity and flight path angle (γ) are unknown functions of time. However, these losses cannot be neglected. Thus, they are estimated based on the current state of the art.

Drag ∆V

As shown in Fig. 4.5, the aerodynamic drag is directed opposite to the velocity and it is given by: 1 D = ρv2SC (4.9) 2 D

From this force, it would be possible to obtain the drag loss applying Eq. 4.8. For the development of this project, it is considered the data provided by Atlas User 1 2 Manual[9], where the dynamic pressure ( 2 ρv ) is represented in front of time.

Figure 4.6: Typical Atlas V 521 Figure 4.7: First order approxi- Dynamic Pressure vs Time[9] 1 2 mation ( 2 ρv vs t) 36 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

In order to get an approximation of the effect of drag, it is used Fig. 4.7, where it is shown that the maximum dynamic pressure is 700 psf (33,500 N/m2) and it is the same for 50 seconds approximately. According to [60], the CD in the atmosphere is of the order of 1. To estimate the reference area (S), it is consid- ered that the payload fairing is a cone for the third part of its length, and this is the surface that is in contact with the air. Regarding the mass, it is considered a value of 400,000 kg, however, this value would be different for each launch vehicle.

Launch Vehicle Diameter [m] Length [m] Reference Area [m2]

Falcon 9[61] 5.2 13.9 43.4 Atlas V[9] 5.4 26.5 78.3 Vega[62] 2.6 7.88 11.9 Delta IV [22] 5.08 19.8 56.4

Average 47

Table 4.7: Payload Fairing Dimensions

Finally, the drag loss can be estimated by:

q · C · S 33, 500 · 1 · 47 ∆V = D · t = · 50 ≈ 200 m/s D m 400, 000

Gravity ∆V

Integrating Eq. 4.8 for a short period of time, it can be expressed by:

∆VG = g0 t sin γ (4.10)

Taking into account that γ is not constant throughout the flight (tb), the gravity loss can be approximated as:

∆VG ≈ g0 tb sin γ (4.11)

According to [60], this sin γ represents the time-averaged value of sin γ. Launch vehicles are launched upward (γ = 90o) and it turns 0o when the launch vehicle reaches the circular orbit. To minimise gravity losses, launch vehicles pitch as soon 37 38 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

as possible, flying more time at lower values of γ. For this project, it is considered that 25o is a reasonable value.

To estimate the time during which the launch vehicle is undergone to gravity effects

(tb), it is based on the current state of the art, where a lot of launch vehicles take between 300 and 450 seconds until they follow a horizontal path. Thus, for the

development of this project, it is considered that tb = 375 seconds.

Finally, gravity loss can be approximated to:

∆VG = 9.81 · 375 · sin 25 = 1, 550 m/s

Both ∆VD and ∆VD have reasonable values compared to other launch vehicles.

Figure 4.8: Launch System Performance Losses[6]

Steering ∆V

As said, the flight path angle varies along the time, causing some losses that also need to be taken into consideration. In this case, it is considered that the steer- ing losses and other losses are 4% of the total Delta-V. Comparing to the Space Shuttle (Table 4.8), whose steering loss is about 4.5% of the ideal Delta-V, this is a reasonable assumption.

38 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

Ideal satellite velocity 7,790 m/s ∆V to overcome gravity losses 1220 m/s ∆V to turn the flight path from the vertical 360 m/s ∆V to counteract aerodynamic drag 118 m/s Orbit injection 145 m/s Deorbit manoeuvre to re-enter atmosphere and aerodynamic braking 60 m/s Correction manoeuvres and velocity adjustments 62 m/s Initial velocity provided by the earth’s rotation at 28.5o latitude -408 m/s

Total required mission velocity 9,347 m/s

Table 4.8: Space Shuttle Incremental Flight Velocity Breakdown[10]

4.3.2 Rocket Equations

The mass of the launch vehicle can be written as the sum of 3 primary components:

• mPL: Payload Mass

• mS: Structure Mass

• mP : Propellant Mass

Thus, at lift-off, the launch vehicle’s mass is:

m0 = mPL + mS + mP (4.12)

Whereas in the end, the launch vehicle has consumed all the propellant and its mass is:

mf = mPL + mS (4.13)

Rockets use different stages (multistage rockets), which allows carrying more pay- load to space than single stage rockets. The main idea of multistage is to discard empty tanks and extra structure. Each stage has its engines and accelerates the payload before being detached.

Both single-stage and multistage rockets satisfy Tsiolkowsky Equation: 39 40 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

m0,i m0,i ∆Vi = g0Isp, i ln( ) or ∆Vi = g0ci ln( ) (4.14) mf,i mf,i

The subscript i represents each stage (i.e. i = 1, 2, 3,..., number of stages) and the total velocity of the launch vehicle must satisfy the following equation:

n X ∆V = ∆Vi (4.15) i=1

To define a launch vehicle, other parameters are also used and must be known:

• Mass Ratio −→ R = m0 mf

• Payload Ratio −→ λ = mPL mP +mS

• Structural Coefficient −→ ε = mS mP +mS

• Propellant Ratio −→ ζ = mP mP +mS

According to Sutton, the mass ratio is the final vehicle mass divided by the initial vehicle mass: mf MR = (4.16) m0

A representation of the various masses is shown in Fig. 4.9:

Figure 4.9: Definitions of Various Vehicle Masses[10]

40 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

4.3.3 Optimal Staging

The optimisation of a launch vehicle aims to carry as much payload possible with the minimum lift-off mass. To do that, the different stages would have different features and not necessarily must be similar among them. Now, stages are defined by the specific impulse (Isp) and the structural coefficient (ε/ kS).

The specific impulse is function of pressure ratio, specific heat ratio, combustion temperature, mixture ratio, and molecular mass, however, in this project it is considered function of the propellant. There are different types of propellants, but for launch vehicle boosters, strap-on engines and upper stages of the launch vehicle, the most used are chemical propellants. As shown in Table 4.9, they are divided into solid and liquid propellants:

Type Propellant Advantages Disadvantages

Limited performance, higher thrust, Solid Simple, reliable, relatively low cost safety issues Liquid: Low performance, higher mass Monopropellant H2O2, N2H4 Simple, reliable, low cost than bipropellant LOX/RP − 1 High performance More complicated system

LOX/LH2 Very high performance Cryogenic, complicated N O /UDMH Storable, good performance Complicated Bipropellant 2 4 F2/N2H4 Very high performance Toxic, dangerous, complicated

OF2/B2H6 Very high performance Toxic, dangerous, complicated

CIF5/N2H4 High performance Toxic, dangerous

Dual Mode N2O4/N2H4 High performance Toxic, dangerous

Table 4.9: Advantages and Disadvantages of Chemical Propellants[6]

The aim of this project is to define the optimal rockets, but always keeping in mind safety aspects. Therefore, the only propellants considered for further calculations are solid propellants, LOX/RP − 1, LOX/LH2 and N2O4/UDMH, which are also the most commonly used by current operating launch vehicles. The reference values of the specific impulse for each propellant are shown below:

41 42 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

Propellant Isp SL [s] Isp Vacuum [s] Solid 260 280-300

N2O4/UDMH 285 300-340 LOX/RP − 1 290 350

LOX/LH2 390 450

Table 4.10: Performance of Chemical Propellants[10]. *These are reference values.

As can be seen in Table 4.10, according to the propellant, the specific impulse is higher or lower. According to [6], to maximise the payload fraction, the stages with higher Isp should be above stages with lower Isp, however, not all launch vehicles satisfy it.

On the one hand, it is necessary to differentiate between sea level and vacuum. The maximum thrust for a given nozzle is obtained when the atmospheric pressure is zero (p3 = 0, vacuum), whereas, between sea level and the vacuum, the thrust varies with the altitude according to Eq. 4.17:

T =mv ˙ 2 + (p2 − p3)A2 (4.17)

All these parameters are shown below:

Figure 4.10: Chamber and Nozzle Representation [10]

42 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

The specific impulse for constant propellant mass flow (m˙ ), and constant thrust, can be written as:

T I = (4.18) sp m˙

Thus, in this study, it is considered that the specific impulse at sea level is used for the first stages, whereas, for upper stages the atmospheric pressure is negligible.

Figure 4.11: Altitude performance of RS 27 liquid propellant rocket engine used in early versions of the Delta launch vehicle [10]

On the other hand, the launch vehicle’s structure is the contribution of the fuel tanks and engines. To obtain a more realistic behaviour of the structure, as far as possible, the structural mass is separated into a contribution from the engine (proportional to the thrust) and a contribution from the tanks (proportional to the propellant). Thus, the new structural coefficient is given by:

a0/g kt (1 − MR) + ε = T/Weng (4.19) (1 + k )(1 − M ) + a0/g t R T/Weng

The term T/Weng is called the thrust to weight ratio. For the development of this thesis, it is estimated based on the current state of the art of rocket engines. The thrust to weight ratio for the first stage is shown in Table 4.11:

43 44 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

Engine Thrust to Weight Ratio

RD-191 89 RD-264 128 RD-181 79 RS-68A 54 Rutherford SL 71

Average 84

Table 4.11: First Stage Engines[18]

For the upper stages it is shown in Table 4.12:

Engine Thrust to Weight Ratio

LE-5B 49 RL10B-2 37 14D30 21 RL10A-4-2 5.3

Average 28

Table 4.12: Upper Stages Engines[19]

On the other hand, kt represents the contribution from the tanks. The higher kt, the higher the lift-off mass. Within this project its values lie between 0.08 and 0.1.

To optimise the mass, it is used the Lagrange multiplier method, which is explained by Howard Curtis [Orbital Mechanics for Engineering Students, p.570 ]. Applying this method, it is obtained a non-linear equation, which is solved by the Newton- Raphson method. To compute all the parameters, a Matlab script (Annex D) is developed using the formulas and the flow chart presented below:

44 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

Structural coefficients:

k (1 − M ) + a0/g t,i R,i T/Weng kS,i = (4.20) (1 + k )(1 − M ) + a0/g t,i R,i T/Weng

Initial Lagrange multiplier guess and equations for optimal staging:

β µ0 = max( ) (4.21) ci(β − kS,i)

n X µcikS,i g(µ) = ∆V + c ln( ) (4.22) i µc − 1 i=1 i n X ci g0(µ) = − (4.23) µ(µc − 1) i=1 i 1 µmin = max( ) (4.24) ci Mass ratios: µcikS,i MR,i = (4.25) µci − 1 Delta-V:

∆Vi = −ci ln(MR,i) (4.26)

Payload ratios: MR,1 − kS,i λi = (4.27) 1 − kS,i Initial masses: mP L,i m0,i = (4.28) λi Propellant masses

mP,i = (1 − MR,i)m0,i (4.29)

Structural masses: kS,i mS,i = mP,i (4.30) 1 − kS,i

45 46 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

Initial guess Start MRguess

Compute Struc-

tural Coefficient kS

Initial guess µ0

Compute Compute: function g(µ) • Mass Ratios (MRc)

• Delta-V Yes |g(µ)|

• Initial Masses

No • Propellant Masses Compute derivative g’(µ) • Structural Masses

Yes Compute search direc- tion dµ = - g(µ)/g’(µ)

No |M - M |

Update Solution

µk+1 = µk + α · dµ

Yes No µk+1>µmin α = α/2

STOP

46 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

This Matlab script computes the lift-off mass according to the all input data. However, the required thrust is not computed, but according to [10], the thrust-to- weight ratio at lift-off (T/m0g0) has values between 1.2 and 2.2 for large surface- launched vehicles. However, within this study, it is covered a narrower frame, between 1.2 and 1.5. Therefore, thrust can be approximated as:

T = (1.2 ÷ 1.5) m0 g0 (4.31)

As explained in Sec. 4.3.3, for the development of this project, the specific impulse only depends on the propellant. For a given propellant, it also depends on other parameters, such as the mixture ratio. However, this analysis is out of the scope of this project. For the different propellants, it is considered average values (reference values) of the specific impulse with a certain margin (± 5 s), which are shown in Table 4.10.

Once all results are obtained through the Matlab script, an average of the lift- off mass is done considering each specific impulse within the specified margin, according to the propellant. In this way, it is possible to approximate both, mass and minimum necessary thrust at lift-off. In Fig. 4.12 it can be seen how the lift-off mass of a two stages rocket varies in function of the specific impulse. This process is also used for larger rockets that have more than two stages.

Figure 4.12: Lift-off Mass for a Two Stages Rocket to Reach 200 km Circular Orbit (i = 51.6o) with 1,000 kg of Payload

47 48 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

4.3.3.1 Results and Algorithm Validation

Before proceeding with further analysis, it is necessary to verify and validate the results. To do that, it is used the graphical method to obtain the optimal staging and later to compare both methods, the graphical and the Lagrange multiplier, for given missions performed by Falcon 9, Rockot and Minotaur I. Even though man- ufacturers do not provide a lot of data about rockets’ performance, it is possible to find some in the User Manuals.

Two Stages - Falcon 9

In this case, the mission consists of reaching a circular orbit, which is 200 km above Earth’s surface and 51.6o of inclination, with a payload of 9,823 kg[61]. Using both methods, they must compute that the take off mass is about 549,000 kg. Results are presented below:

Figure 4.13: Optimal Staging with Graphical Method (Falcon 9)

48 4.3. PROPULSION SYSTEM’S PARAMETERS CALCULATION ESEIAAT

∆V1/∆V m01 [kg] mp1 [kg] ms1 [kg] ∆V2/∆V m02 [kg] mp2 [kg] ms2 [kg] Graphical 0.462 545,800 436,650 29,021 0.538 79,703 65,097 4,783 Method Lagrange Multiplier 0.461 545,800 436,460 29,010 0.539 80,333 65,684 4,826 Method

Table 4.13: Comparison between the Graphical Method and the Lagrange Multiplier Method (Falcon 9)

Three Stages - Rockot

According to Rockot’s User Manual, this launch vehicle can place up to 2,140 kg of payload into 200 km circular orbit inclined at 63.2o. Using the same methods as for the Falcon 9, but taking into consideration that now the launch vehicle has three stages, it should be obtained that the take-off mass is 107,000 kg. Results for both methods can be seen below:

Figure 4.14: Optimal Staging with Graphical Method (Rockot)

49 50 CHAPTER 4. STUDY OF PROPULSION SYSTEM’S REQUIREMENTS

∆V1/∆V m01 [kg] ∆V2/∆V m02 [kg] ∆V3/∆V m03 [kg] Graphical 0.24 103,780 0.38 38,436 0.38 9,016 Method Lagrange Multiplier 0.235 103,810 0.381 39,341 0.384 9,186 Method

Table 4.14: Comparison between the Graphical Method and the Lagrange Multiplier Method (Rockot)

Four Stages - Minotaur

According to [63], Minotaur I, which is a 36,000 kg solid rocket, can place up to 530 kg into 185 km circular orbit inclined at 70o from PSCA spaceport. In this case, the graphical method was not applied. On the other hand, the Lagrange Multiplier Method was applied in the same way as in the examples above.

Real Mass [kg] 36,200 Computed Mass [kg] 36,027

Relative Error (%) 0.48

Table 4.15: Comparison between Real Mass and Computed Mass (Minotaur-I)

Comparing both methods, it can be seen that results are almost identical, and the difference between the real take-off mass and the computed is almost negligible. Therefore, the Lagrange multiplier method has been applied correctly, and results can be considered reliable.

50 Chapter 5

Results Analysis

In this chapter, it is done the comparison between the results obtained applying the Lagrange Multiplier method and current launch vehicles, in order to select which are the optimal rockets for each mission. In the end, it is studied if the idea of the Rocket launching Consultancy is feasible based on the hypothesis made in Sec. 1.2.

5.1 Results Comparison

In this section are presented only the most optimal launch vehicles for different missions. In Annex C it can be seen all the procedures used to obtain these results, and all the rockets that have been considered.

51 52 CHAPTER 5. RESULTS ANALYSIS

5.1.1 Polar Orbit

1,000 kg 500 kg 100 kg 10 kg

LV Beta Alpha Electron Two Stages LV Thrust / Required Thrust [kN] 2,208 / 1,271-1,589 736 / 636-795 162 / 127-159 LV Mass / Required Mass [kg] 149,700 / 108,000 54,000 / 54,000 10,500 / 10,800

LV Three Stages LV Thrust / Required Thrust [kN] LV Mass / Required Mass [kg]

LV Vega Four Stages LV Thrust / Required Thrust [kN] 2,260 / 1,161-1,451 LV Mass / Required Mass [kg] 133,770 / 98,600

LV Minotaur V Five Stages LV Thrust / Required Thrust [kN] 1,607 / 1,024-1,280 LV Mass / Required Mass [kg] 89373 / 87000

Table 5.1: Optimal Launch Vehicles to Reach Polar Orbit according to the Mission

5.1.2 Sun-Synchronous Orbit

1,000 kg 500 kg 100 kg 10 kg

LV Alpha Electron Vector-H Two Stages LV Thrust / Required Thrust [kN] 736 / 536-670 162 / 107-134 290 / 107-134 LV Mass / Required Mass [kg] 54,000 / 45,500 10,500 / 9,100 8,700 / 9,100

LV Three Stages LV Thrust / Required Thrust [kN] LV Mass / Required Mass [kg]

LV Four Stages LV Thrust / Required Thrust [kN] LV Mass / Required Mass [kg]

LV Minotaur IV Five Stages LV Thrust / Required Thrust [kN] 1,607 / 908 -1,135 LV Mass / Required Mass [kg] 86,300 / 77,000

Table 5.2: Optimal Launch Vehicles to Reach Sun-Synchronous Orbit accord- ing to the Mission

52 5.1. RESULTS COMPARISON ESEIAAT

5.1.3 Inclined Orbit

1,000 kg 500 kg 100 kg 10 kg

LV Vector-R Two Stages LV Thrust / Required Thrust [kN] 89 / 77-96 LV Mass / Required Mass [kg] 6,000 / 6,500

LV Safir Three Stages LV Thrust / Required Thrust [kN] 334 / 327-408 LV Mass / Required Mass [kg] 26,000 / 27,750

LV Minotaur I Four Stages LV Thrust / Required Thrust [kN] 935 / 396-495 LV Mass / Required Mass [kg] 36,200 / 33,650

LV Five Stages LV Thrust / Required Thrust [kN] LV Mass / Required Mass [kg]

Table 5.3: Optimal Launch Vehicles to Reach Inclined Orbit according to the Mission

53 54 CHAPTER 5. RESULTS ANALYSIS Vega C Hyperbola I Kuaizhou 1 Soyuz 2.1v 1,000 kg 500 kg 100 kg 10 kg Optimal Launch Vehicles to Reach Semi-Synchronous Orbit according to the Mission Table 5.4: LVLV Thrust / Required ThrustLV [kN] Mass / Required 3,256 Mass + [kg] 2,962 / 4,862-6,077 3,844 + 1,922 / 4,203-5,253 464,000 / 413,000 481,000 / 357,000 LM2F 1,922 / 2,001 - 2,502 Angara 3 171,000 / 170,000 Angara 1.2 LV Thrust / Required ThrustLV [kN] Mass / Required Mass [kg]LV LV LV Thrust / Required ThrustLV [kN] Mass / Required Mass [kg] 1,920 / 1,371-1,714 157,000 / 116,500 LV LV Thrust / Required ThrustLV [kN] Mass / Required Mass [kg] 4,323 / 1,854-2,318 770 / 473-592 220,000 / N/A 157,500 / 371-464 42,000 / 40,200 30,000 / 31,500 Two Stages Five Stages Four Stages Three Stages 5.1.4 Semi-Synchronous Orbit

54 5.1. RESULTS COMPARISON ESEIAAT Vector-R 89 / 54-67 6,000 / 4,580 1,000 kg 500 kg 100 kg 10 kg Optimal Launch Vehicles to Reach Molnyia Orbit according to the Mission Table 5.5: LVLV Thrust / Required ThrustLV [kN] Mass / Required Mass [kg]LV LV Thrust / Required ThrustLV [kN] Mass 4,323 / / Required 2,413-3017 Mass [kg] 220,000 / 205,000 2,260 / 1,401-1,751 N/A / 1,207-1,508 133,700 / 119,000 120,000 / Vega 102,500 C Vega SSLV LVLV Thrust / Required ThrustLV [kN] Mass / Required 1,920 Mass / [kg] 1,642-1,854 2,962 / 2,413-3,017 157,000 / 139,500 233,000 / 205,000 Soyuz 2.1v LM2C 334 / 241-302 26,000 / 20,500 Safir LV LV Thrust / Required ThrustLV [kN] Mass / Required Mass [kg] Two Stages Five Stages Four Stages Three Stages 5.1.5 Molnyia Orbit

55 56 CHAPTER 5. RESULTS ANALYSIS

Once chosen the optimal launch vehicles for each mission, it is possible to draw some conclusions. But first, let’s make a simple classification of the missions.

Low ∆V (<10,500 m/s) High ∆V (>10,500 m/s) Medium Payload (approx. 1,000 kg) Low Requirements Missions Medium Requirements Missions Low Payload (<100 kg) Low Requirements Missions Low Requirements Missions

Table 5.6: Missions Classification

The main conclusion is that there are not a lot of optimal launch vehicles, neither for medium requirements missions nor low requirements missions. However, for medium requirements missions, there are more than for the low ones. Even so, there are not more than two optimal launch vehicles for those missions.

As has been said, each propellant has its pros and cons. Regarding the use of liquid oxygen, it is not corrosive, however, it is difficult to store. It can be combined with liquid hydrogen, which requires large vehicle volumes and suitable materials for the tanks and the pipes, thus, these types of launch vehicles would be more expensive. On the other hand, using RP-1 instead of liquid hydrogen, it would make the launch vehicle a bit cheaper, because hydrocarbon fuels are easier to handle and they are available at low cost. Therefore, this type of propellant would be a great option that combines, both performance and costs.

Regarding the use of nitrogen tetroxide (N2O4), it can be stored relatively easy in combination with UDMH, which is a fairly stable liquid. It could be an interesting option that provides good performance at a reasonable cost. However, it must be aware that it can be toxic. Last, but not least, solid rockets cannot perform large missions with high performance. However, due to their simple design and easiness of operation, they could be a good option for low requirements missions at a relatively low cost.

Analysing the features of the optimal launch vehicles, it can be seen that for low requirements missions, they are solid four or five stages rockets or LOX/RP − 1 two stages rockets. Whereas for medium requirements missions, optimal rockets have different features. These are two and three stages launch vehicles that use

LOX/RP − 1, N2O4/UDMH or a combination of both, and also four stages solid rockets. On the other hand, LOX/LH2 launch vehicles are too large for these types of missions, and they are orientated towards larger missions with larger budgets

56 5.2. FEASIBILITY STUDY ESEIAAT that require higher performance. Moreover, there are other launch vehicles that use boosters, but in the same way as the previous one, they probably would be optimal for larger missions, which are beyond the scope of this project.

One of the main aspects to consider at booking a launch vehicle is the price. This varies depending on the provider and rocket’s features, for example, storable and non-toxic propellants are cheaper than those that are toxic or cryogenic because there are fewer concerns to take care of regarding safety, storage ,or design. An- other aspect that has a direct impact on the cost is the number of stages, the more stages, the more expensive is the launch vehicle. Therefore, selecting rockets with fewer stages probably would be cheaper. However, this would depend on the launch cost set by the manufacturer, whose study is beyond the scope of this project.

5.2 Feasibility Study

The main hypothesis to determine the feasibility of this type of business is the number of optimal launch vehicles for each mission, being feasible if there are not many.

Analysing all the missions, there are not more than two optimal rockets for each one. Therefore, booking a launch vehicle that satisfies the necessities of mission and it is not oversized can be more difficult than expected. Even for some low requirements missions, there is only a single optimal launch vehicle.

This is where the consultancy becomes important, because selecting the optimal launch vehicle amongst a few requires time and knowledge. The consultancy would provide this service at an affordable price, making the customers not worrying about the planning phase.

For this type of business, it is very important to have a contract with rocket manufacturers, in order to get all the information possible and also to be able to contact them easier. In Sec. 5.1, it is seen that for low requirements missions there is hardly any optimal launch vehicle. However, there are start-ups that own or are designing launch vehicle for micro-launchers (usually low payload to LEO).

57 58 CHAPTER 5. RESULTS ANALYSIS

At the start of this business, being an unknown company, it would be difficult to get contracts with well-known rocket manufacturers and launch services providers. However, with start-ups, which are neither well-known, it would be possible to reach an agreement where both could be benefited. On the one hand, the start-up would have more possibilities to be hired to perform a launch. On the other hand, the consultancy would be able to increase its catalogue. And in the future, trying to get as many contracts as possible, to offer more options to the customers that hire our services. Some of these new start-ups are listed below:

• Pangea Aerospace

• PLD Space

• OneSpace

• Vector

Taking into consideration these aspects, and also that there is no direct competitor except Precious Payload, which is also a new company (2017), it would be easier to get this market niche and making this idea a profitable business.

5.2.1 Other Services

As stated, current launch vehicles are mainly focused on large missions. Whereas for low requirements missions there are only a few optimal rockets, although, there are new start-ups that are designing rockets for micro-launchers to LEO. Despite the apparition of these new start-ups, their rockets would not be available until the next years.

In order to optimise the current launch vehicles for low requirements missions, some manufacturers provide something called "Ride-Share", which consists of sharing a launch, making the launch more affordable.

Apart from taking care of the planning phase, this new idea of Ride-Share could allow the consultancy to provide another service to the customers. This would consist of having a platform where launch services providers could upload their offers of Ride-Share and the customers could book a launch if they were interested. 58 5.2. FEASIBILITY STUDY ESEIAAT

This would consist as follows. On the one hand, the manufacturers should upload the requirements of the mission, which are payload’s features and the target orbit. It would also be necessary an approximate date of the launch. On the other hand, the customers would be able to sort according to their necessities, and the consultancy would indicate through this platform, which launches satisfy better their necessities.

This is service is also provided by other companies, for example, Spaceflight Inc. or The Precious Payload. However, providing this service would allow the company to have a higher market share.

59

Chapter 6

Conclusions and Future Work

This project aims, on the one hand, to obtain the optimal rockets for a few missions through the analysis of the propulsion system’s requirements. On the other hand, studying the feasibility of a "Rocket launching Consultancy" based on the number of optimal launch vehicles for each mission.

To successfully achieve the objectives, it was necessary to obtain the optimal launch vehicles for each mission. To do this, Lagrange Multiplier Method and Tsi- olkovsky’s equation (Eq. 4.14) were applied to obtain the minimum take-off mass according to the mission and the characteristics of the launch vehicle (propellant type and structural coefficient).

Regarding optimal launch vehicles, it was obtained that for low requirements mis- sions, the optimal rockets are those that have two stages and use LOX/RP − 1 and solid four stages rockets. Whereas for medium requirements missions, the optimal rockets are those that have two or three stages and use LOX/RP − 1 and N2O4/UDMH. Last but not least, LOX/LH2 rockets are oversized for the missions studied within this project and are likely designed for larger missions, for example, to reach other planets or carrying heavier payloads.

Considering there are not many optimal rockets, and currently, there is almost no company that provides similar services, the idea of the consultancy proposed within this project it is considered to be feasible. The services provided would

61 62 CHAPTER 6. CONCLUSIONS AND FUTURE WORK be guidance throughout the planning phase and a platform where to book a ride- share. All this at an affordable price for new businesses, to make space more affordable for everyone.

It must be taken into account that this is an academic project, which has its limitations. Each launch vehicle is different, however, the calculations were done with reference values for the specific impulse and structural coefficient. Regardless of not studying each rocket in detail, it was possible to get an idea that current rockets focus on larger missions, not in low requirements missions. Furthermore, considering that the small satellite market is growing and there are not many small rockets, it might be interesting for rocket manufacturers to design rockets that focus on micro-launchers.

In conclusion, all objectives have been achieved, and it was obtained that there are not many optimal rockets for a specific mission. Therefore, a business that provides the services mentioned within this project could become feasible.

Future Work

This project studied the feasibility of this type of business, but there more tasks that can be done to become this idea into a reality:

• Development of a Business Plan: this would allow to know better the initial investment and how much it would take to recover it.

• Study of larger missions: it would make to compare different missions and it would be easier to decide in what market niche is more feasible to have benefits.

• Detailed study of each launch vehicle: with the specific values of the specific impulse and the structural coefficient of each launch vehicle it would be possible to know their performance, not only an approximation. This also could be used to develop a tool that would make the selection of the optimal rockets easier according to the mission. This would also entail more validation of the results.

62 Environmental Impact

Nowadays, rockets are the only way to place a payload into orbit, to explore our Solar System and beyond. It will remain in this way in the years to come due to the limitations of the current technology. Nowadays, most of the mass of a rocket is propellant, which is up to 85-90% of its mass.

The idea of the consultancy is to choose the optimal rockets for the different mis- sions, and indirectly making possible to reduce the propellant used. Depending on the propellant, rockets produce different emissions, which are mainly carbon diox- ide (CO2), water vapour, carbon soot, carbon monoxide, NOx, chlorine, alumina, and sulphuric compounds. It must be taken into account that not all propellants produce the same amount of emissions. In Table 6.1 it can be seen some examples of rockets’ emissions and a comparison with aviation:

Water Sulphuric Rocket Fuel CO2 Soot NOx Chlorine Alumina Vapour Compounds

Titan II Hypergolic 36 16 0.2 0.3 0 0 0.3 RP-1 + Soyuz FG 243 64 13 0.4 0 0 ∼0 Hypergolic Falcon 9 RP-1 425 152 30 1 0 0 ∼0 Delta IV Heavy Hydrogen ∼0 632 0 0.5 0 0 0 SRB + Space Shuttle 443 976 4.2 7 250 350 ∼0 Hydrogen Boeing 747 Kerosene 302 N/A N/A N/A N/A N/A N/A Boeing 737 Kerosene 60 N/A N/A N/A N/A N/A N/A

Table 6.1: Rocket and Aviation Emissions[Metric T onnes] [20]

Throughout this business, it is intended to avoid that customers book oversized launch vehicles for their spacecraft, in order to reduce costs and environmental impact.

63 64 CHAPTER 6. CONCLUSIONS AND FUTURE WORK

Regarding the development of this project, it has been done only using a computer. Considering that it has a power of 500 W and it has been used for 600 hours, the total energy used can be seen below:

T otal Energy = 0.5 kW · 600 h = 300 kW h

According to [64] 1 kWh of electricity produces 200 g of CO2 and 0.52 mg of radioactive wastes. Therefore, for the development of this project it has been produced approximately the following amounts of carbon dioxide and radioactive wastes:

kg CO CO Emissions = 300 kW h · 0.2 2 = 60 kg CO 2 kW h 2

mg RW Radioactive W astes (RW ) = 300 kW h · 0.52 = 156 mg RW kW h

Another aspect to take into consideration is the emissions due to the use of a vehicle. There were only two meetings at the university because of the outbreak of COVID-19. Considering that the average distance between Esparreguera and

Terrassa is 25 km and the average CO2 emissions the car used to travel is 138 g/km according to the manufacturer [65]. The total emissions due to the use of a car are the following:

CO2 Emissions = 4 · 25 km · 138 g/km = 13.8 kg of CO2

64 Budget

In order to obtain a more realistic budget, it has been divided into different groups. These are human resources, software licenses, energy, and equipment.

Human Resources

This is the author’s fees, who earns 18 e/hour. The duration of this project is 600 hours (see Planning for more detail):

18 e Human Resources = · 600 hours = 10, 800 e hour

Software Licenses

For the development of this project it was necessary some software, which is listed below:

Software Licenses Units Cost/Unit Cost

Matlab 1 800.00 e 800.00 e Microsoft Office 1 69.00 e 69.00 e

Total 869.00 e

Table 6.2: Software Licenses Budget

65 66 CHAPTER 6. CONCLUSIONS AND FUTURE WORK

Energy and Equipment

Here are included the cost of the electricity, fuel and all materials needed for the development of this project:

Energy and Equipment Units Cost/Unit Cost

Electricity [kWh] 300 0.1586 e/kWh 47.58 e Fuel [L] 5.5 1.40 e/L 7.70 e Office Supplies 1 50.00 e 50.00 e Personal Computer 1 1,000.00 e 1,000.00 e

Total 1,105.28 e

Table 6.3: Energy and Equipment Budget

Total Budget

The total budget of this project is shown below:

Expenses Unit Cost/Unit Cost Total Cost

Human Resources Author’s fees 600 h 18.00 e/h 10,800 e 10,800 e

Matlab 1 800.00 e 800.00 e Software Licenses 869.00 e Microsoft Office 1 69.00 e 69.00 e

Electricity 300 kWh 0.1586 e/kWh 47.58 e Fuel 5.5 L 1.40 e/L 7.70 e Energy and Equipment 1,105.28 e Office Supplies 1 50.00 e 50.00 e Personal Computer 1 1,000.00 e 1,000.00 e

Total Budget 12,774.28 e

Table 6.4: Total Budget

66 Planning

In order to carry out this study and to comply with the scope, which is defined at the beginning of this document, the planning of the tasks is required. Otherwise, it would not be possible to have control over the developed work throughout this project, making impossible to achieve its aims.

The planning was presented in the Project Charter, and it was followed as much as possible. All the tasks developed in this thesis are listed in Table 6.5. In order to do a sort of quality control, the hours spend on each task have been taken into account, and they can be seen in Table 6.6.

Finally, how this project was developed can be seen in the Gantt Diagram. This is quite similar to the presented in the Project Charter, with some differences.

67 68 CHAPTER 6. CONCLUSIONS AND FUTURE WORK

Task Code Task Description Preceding Task 0 Introduction 0.1 Brief General Research - 0.2 Business Description - 0.3 Brief Market Research 0.2 1 State of the art 1.1 Payload Research - 1.2 Launching Sites Research - 1.3 Orbits Research - 1.4 Operating Rockets Research - 2 Study of Propulsion System’s Requirements 2.1 Missions definition 1.1, 1.2, 1.3 2.2 Study of each mission 2.2.1 Study of the possibility to reach the orbit 2.1 2.2.2 Propulsion System’s Parameters Calculation 2.2.1 3 Result Analysis 3.1 Algorithm Validation 2.2.2 3.2 Results Comparison 3.1 3.3 Feasibility Study 3.2 4 Deliverable 4.1 Project Charter - 4.2 First Follow-up - 4.3 Second Follow-up - 4.4 Third Follow-up - 5 Report Development 6 Final Delivery 6.1 Report Revision 5 6.2 Budget 5 6.3 Annex 5 7 Presentation 7.1 Presentation Development 6 7.2 Thesis Presentation and Defence 7.1

Table 6.5: Bachelor’s Thesis Tasks

68 ESEIAAT

Task Code Task Description Level of effort (hours) 0 Introduction 40 0.1 Brief General Research 10 0.2 Business Description 10 0.3 Brief Market Research 20 1 State of the Art 50 1.1 Payload Research 10 1.2 Launching Sites Research 10 1.3 Orbits Research 10 1.4 Operating Rockets Research 20 2 Study of Propulsion System’s Requirements 225 2.1 Missions definition 20 2.2 Study of each mission (205) 2.2.1 Study of the possibility to reach the orbit 40 2.2.2 Study of Propulsion system’s parameters 165 3 Result Analysis 95 3.1 Algorithm Validation 25 3.2 Results Comparison 40 3.3 Feasibility Study 30 4 Deliverable 35 4.1 Project Charter 32 4.2 First Follow-up 1 4.3 Second Follow-up 1 4.4 Third Follow-up 1 5 Report Development 90 6 Final Delivery 65 6.1 Report Revision 35 6.2 Budget 10 6.3 Annex 20 7 Presentation - 7.1 Presentation Development - 7.2 Thesis Presentation and Defence - Total Effort 600

Table 6.6: Level of effort to develop each task

69 70 CHAPTER 6. CONCLUSIONS AND FUTURE WORK

Figure 6.1: Gantt Diagram (1/2)

70 ESEIAAT

Figure 6.2: Gantt Diagram (2/2)

71

Appendix A

Operating Launch Vehicles

A.1 Two Stage Launch Vehicles

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Mass [kg] Launch Site Propellant Type Propellant Type 736 in vacuum 70 CCAFS Alpha[66] 54,000 LOX/RP − 1 LOX/RP − 1 VAFB 2,208 in vacuum 163 CCAFS Beta[66] N/A LOX/RP − 1 LOX/RP − 1 VAFB 162 22 Electron[53] 10,500 Mahía LOX/RP − 1 LOX/RP − 1 CCAFS 7,607 934 Falcon 9 FT 549,054 KSC LOX/RP − 1 LOX/RP − 1 VAFB 22,819 934 KSC Falcon Heavy 1,420,788 LOX/RP − 1 LOX/RP − 1 VAFB 2,961.6 742 Long March 2D[67] 232,250 Jiuquan N2O4/UDMH N2O4/UDMH 290 13 CCAFS Vector-H[68] 8,700 LOX/RP − 1 LOX/RP − 1 VAFB 89 4.5 CCAFS Vector-R[68] 6,000 LOX/RP − 1 LOX/RP − 1 VAFB

Table A.1: Two Stage Launch Vehicles without Boosters Configuration [21]

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Total Boosters Thrust [kN] Mass [kg] Launch Site Propellant Type Propellant Type 960 67 Ariane 5[69] 14,000 780,000 GSC LOX/LH2 LOX/LH2 2,962 831 Long March 2F[67] 3,256 464,000 Jiuquan N2O4/UDMH N2O4/UDMH 1,598 186 LVM3[23] 10,300 640,000 Satish Dhawab N2O4/UDMH LOX/LH2

Table A.2: Two Stage Launch Vehicles with Boosters Configuration [21]

73 74 APPENDIX A. OPERATING LAUNCH VEHICLES

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Total Boosters Thrust [kN] Mass [kg] Launch Site Propellant Type Propellant Type 1,922 294.3 Plesetsk Angara 1.2 - 171,000 LOX/RP-1 LOX/RP − 1 Vostonchy 1,922 294.3 Plesetsk Angara 3 3,844 481,000 LOX/RP − 1 LOX/RP − 1 Vostonchy 1,922 294.3 Plesetsk Angara 5 7,688 773,000 LOX/RP − 1 LOX/RP − 1 Vostonchy

Table A.3: Angara Launch Vehicle Family [21]

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Total Boosters Thrust [kN] Mass [kg] Launch Site Propellant Type Propellant Type 3,827 99.2 or 198.4 CCAFS Atlas V 401 - 333,731 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 411 1,688 approx. 380,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 421 3376 approx. 427,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 431 5,064 approx. 474,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 501 - approx. 315,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 511 1,688 approx. 362,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 521 3,376 approx. 408,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 531 5,064 approx. 455,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 541 6,752 approx. 502,000 LOX/RP − 1 LOX/RP − 1 VAFB 3,827 99.2 or 198.4 CCAFS Atlas V 551 8,440 568,878 LOX/RP − 1 LOX/RP − 1 VAFB

Table A.4: Atlas V Launch Vehicle Family [21][9]

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Total Boosters Thrust [kN] Mass [kg] Launch Site Propellant Type Propellant Type Delta IV 3,137 110 CCAFS - 249,500 Medium LOX/LH2 LOX/LH2 VAFB Delta IV 3,137 110 CCAFS 2,491 or 4,982 - Medium+(4,2) LOX/LH2 LOX/LH2 VAFB Delta IV 3,137 110 CCAFS 2,491 or 4,982 - Medium+(5,2) LOX/LH2 LOX/LH2 VAFB Delta IV 3,137 110 CCAFS 2,491 or 4,982 - Medium+(5,4) LOX/LH2 LOX/LH2 VAFB Delta IV 3,137 110 CCAFS 6,274 733,000 Heavy LOX/LH2 LOX/LH2 VAFB

Table A.5: Delta IV Launch Vehicle Family [21][22]

Thrust 1st Stage at Sea Level [kN] Thrust 2nd Stage [kN] Type Total Boosters Thrust [kN] Mass [kg] Launch Site Propellant Type Propellant Type 1,098 in vacuum 137 H-IIA 202 4,520 289,000 Tanegashima LOX/LH2 LOX/LH2 1,098 in vacuum 137 H-IIA 204 9,040 443,000 Tanegashima LOX/LH2 LOX/LH2 2,196 in vaccum 137 H-IIB 6,320 530,000 Tanegashima LOX/LH2 LOX/LH2

Table A.6: H-II Launch Vehicle Family [21][19]

74 A.2. THREE STAGE LAUNCH VEHICLES ESEIAAT

A.2 Three Stage Launch Vehicles

Thrust 1st Stage at SL [kN] Thrust 2nd Stage [kN] Thrust 3rd Stage [kN] Type Mass [kg] Launch Site Propellant Type Propellant Type Propellant Type 3,265 396.3 77.8 Antares 240,000 MARS LOX/RP − 1 Solid Solid 4,160 755 18.6 Baikonour Dnepr 201,000 N2O4/UDMH N2O4/UDMH N2O4/UDMH Domarovsky Jiuquan 2,961.6 741.3 10.8 Long March 2C[67] 233,000 Taiyuan N2O4/UDMH N2O4/UDMH Solid Xichang 2,962 742 167 Long March 3A[67] 241,000 Xichang N2O4/UDMH N2O4/UDMH LOX/LH2 2,962 741.3 206 Jiuquan Long March 4B[67] 249,000 N2O4/UDMH N2O4/UDMH N2O4/UDMH Taiyuan 2,962 741.3 201.8 Jiuquan Long March 4C 250,000 N2O4/UDMH N2O4/UDMH N2O4/UDMH Taiyuan 1,179 175 16 Long March 6 103,217 Taiyuan LOX/RP − 1 LOX/RP − 1 H2O2/RP-1 10,000 2,400 19.2 Proton Medium 665,000 Baikonour N2O4/UDMH N2O4/UDMH N2O4/UDMH 1,870 240 19.6 Rockot 107,000 Plesetsk N2O4/UDMH N2O4/UDMH N2O4/UDMH 334.4 19.6 - Safir 26,000 Semnan N2O4/UDMH N2O4/UDMH Solid 564 564 60.4 Shavit 2 70,000 Palmachin AFB Solid Solid Solid 1,920 297.9 2.94 Baikonour Soyuz 2.1v 157,000 LOX/RP − 1 LOX/RP − 1 UDMH Plesetsk 1,900 240 4.9 Strela 105,000 Baikonour N2O4/UDMH N2O4/UDMH N2O4/UDMH 1,100 250 54 Unha 90,000 Sohae RFNA/UDMH RFNA/UDMH LOX/RP − 1 7,256 992 79.5 Zenit 470,000 Baikonour LOX/RP − 1 LOX/RP − 1 LOX/RP − 1 440 N/A N/A Zhuque-1 27,000 Jiuquan Solid Solid Solid

Table A.7: Three Stage Launch Vehicles without Boosters Configuration [21]

Total Boosters Thrust 1st Stage at SL [kN] Thrust 2nd Stage [kN] Thrust 3rd Stage [kN] Type Mass [kg] Launch Site Thrust [kN] Propellant Type Propellant Type Propellant Type 4,700 396.3 75 GSLV 2,720 414,750 Satish Dhawan Solid N2O4/UDMH LOX/LH2 2,962 742 167 Long March 3B/E[67] 2,961.6 485,970 Xichang N2O4/UDMH N2O4/UDMH LOX/LH2 2,962 742 167 Long March 3C[67] 1,480 345,000 Xichang N2O4/UDMH N2O4/UDMH N2O4/UDMH 1,018 176.6 78.5 Long March 5 9,432 879,000 Whenchang LOX/LH2 LOX/LH2 LOX/LH2 4,716 2,358 700 Long March 7 Boosters only 594,000 Whenchang LOX/RP − 1 LOX/RP − 1 LOX/RP − 1 838.5 297.9 19.9 Plesetsk Soyuz 2.1a/b[59] 3,170 305,000 LOX/RP − 1 LOX/RP − 1 N2O4/UDMH GSC

Table A.8: Three Stage Launch Vehicles with Boosters Configuration [21]

75 76 APPENDIX A. OPERATING LAUNCH VEHICLES

A.3 Four Stage Launch Vehicles

Thrust 1st Stage Thrust 2nd Thrust 3rd Thrust 4th Type at SL [kN] Stage [kN] Stage [kN] Stage [kN] Mass [kg] Launch Site Propellant Type Propellant Type Propellant Type Propellant Type 1,580 377.2 81.3 <1 Epsilon 90,800 Uchinoura SC Solid Solid Solid Hydrazine 770 597 195 60 Hyperbola-1[70] 42,000 Jiuquan Solid Solid Solid Solid N/A N/A N/A N/A Kuaizhou 1 30,000 Jiuquan Solid Solid Solid N2O4/UDMH N/A N/A N/A N/A Kuaizhou 11 78,000 Jiuquan Solid Solid Solid N2O4/UDMH 935 268 118.2 34.8 MARS Minotaur I 36,200 Solid Solid Solid Solid VAFB CCAFS 1,904 704 36 36 or 47.3 MARS Minotaur-C 77,000 Solid Solid Solid Solid PSCA VAFB 10,000 2,400 583 19.2 Proton M 705,000 Baikonour N2O4/UDMH N2O4/UDMH N2O4/UDMH N2O4/UDMH N/A N/A N/A N/A SSLV 120,000 Satish Dhawan Solid Solid Solid Liquid 2,260 1,120 317 2.45 Vega[62] 133,770 GSC Solid Solid Solid Solid 4,323 1,304 317 2.45 Vega-C[71] 220,000 GSC Solid Solid Solid N2O4/UDMH

Table A.9: Four Stage Launch Vehicles Without Booster Configuration [21]

Thrust 1st Stage Thrust 2nd Thrust 3rd Thrust 4th Total Boosters Type at SL [kN] Stage [kN] Stage [kN] Stage [kN] Mass [kg] Launch Site Thrust [kN] Propellant Type Propellant Type Propellant Type Propellant Type 4,800 799 240 14.6 PSLV-CA - 230,000 Satish Dhawan Solid N2O4/UDMH Solid Solid 4,800 799 240 14.6 PSLV-G 4,314 295,000 Satish Dhawan Solid N2O4/UDMH Solid Solid 4,800 79 240 14.6 PSLV-XL 4,314 320,000 Satish Dhawan Solid N2O4/UDMH Solid Solid

Table A.10: PSLV Launch Vehicle Family [21][23]

76 A.4. FIVE STAGE LAUNCH VEHICLES ESEIAAT

A.4 Five Stage Launch Vehicles

Thrust 1st Stage Thrust 2nd Thrust 3rd Thrust 4th Thrust 5th Type at SL [kN] Stage [kN] Stage [kN] Stage [kN] Stage [kN] Mass [kg] Launch Site Propellant Type Propellant Type Propellant Type Propellant Type Propellant Type CCAFS 1,607 1,365 329 64 or 34.8 34.8 MARS Minotaur IV 86,300 Solid Solid Solid Solid Solid PSCA VAFB CCAFS 1,607 1,365 329 64 47.3 MARS Minotaur V 89,373 Solid Solid Solid Solid Solid PSCA VAFB CCAFS 1,607 1,607 1,365 329 64 or 34.8 MARS Minotaur VI 89,373 Solid Solid Solid Solid Solid PSCA VAFB

Table A.11: Five Stage Launch Vehicles without Booster Configuration [21]

77

Appendix B

Possibility to Reach the Orbits

*Note: Red Box: Those are the sites from where the orbit cannot be reached. Green Box: Those are the launch azimuths to reach the target orbit.

79 80 APPENDIX B. POSSIBILITY TO REACH THE ORBITS

B.1 Polar Orbit

o o o o o Launch Site ψ01 [ ] ψ02 [ ] ψ0 [ ] V0 [m/s] Vf [m/s] ∆Videal [m/s] ψR1 [ ] ψR2 [ ] Plesetsk 16.65 163.35 16.65 211.9 8,126.8 8,068.67 15.21 164.79 Baiknour 10.83 169.17 10.83 323.1 8,126.8 8,072.35 8.58 171.42 CCAFS/KSC 8.55 171.45 - 408.3 8,126.8 8,076.21 5.68 174.32 VAFB 9.12 170.88 170.88 382.8 8,126.8 8,074.96 6.44 173.56 Guiana SC 7.53 172.47 7.53 463.2 8,126.8 8,079.17 4.27 175.73 Xichang 8.53 171.47 - 409.5 8,126.8 8,076.27 5.65 174.35 Jiuquan 9.96 170.04 - 351.0 8,126.8 8,073.52 7.51 172.49 Tanegashima 8.70 171.30 8.70 401.1 8,126.8 8,075.85 5.89 174.11 Taiyuan 9.66 170.34 - 362.0 8,126.8 8,074.00 7.12 172.88 Satish Shawan 7.72 172.28 7.72 451.9 8,126.8 8,078.53 4.54 175.46 MARS 9.52 170.48 - 367.0 8,126.8 8,074.23 6.95 173.05 Uchinoura 8.79 171.21 - 397.4 8,126.8 8,075.67 6.00 174.00 Yasny 12.00 168.00 12.00 292.1 8,126.8 8,071.17 9.97 170.03 Palmachin 8.84 171.16 - 394.8 8,126.8 8,075.54 6.07 173.93 Imam 9.19 170.81 - 380.0 8,126.8 8,074.83 6.53 173.47 Vostochny 12.21 167.79 12.21 287.0 8,126.8 8,070.99 10.22 169.78 Mahia 9.71 170.29 170.29 359.9 8,126.8 8,073.91 7.19 172.81 PSCA 14.02 165.98 165.98 250.6 8,126.8 8,069.78 12.29 167.71 Sohae 9.77 170.23 170.23 357.8 8,126.8 8,073.82 7.26 172.74 Wenchang 7.96 172.04 172.04 438.1 8,126.8 8,077.78 4.88 175.12

Table B.1: Calculations of thePossibility to Reach the Polar Orbit

Manoeuvres

Elliptical Orbit (Perigee = 200 km and Apogee = 1,800 km):

s 1 1 uapogee = 2µE( − ) = 6, 785.8 m/s ra rp + ra

Circular Orbit (Radius = 1,800 km)

rµ u = E = 7, 133.1 m/s circ r

Delta-V needed for manoeuvres:

∆Vman = ucirc − uapogee = 7, 133.1 − 6, 785.8 = 327.4 m/s

80 B.2. SUN-SYNCHRONOUS ORBIT ESEIAAT

B.2 Sun-Synchronous Orbit

o o o o o Launch Site ψ01 [ ] ψ02 [ ] ψ0 [ ] V0 [m/s] Vf [m/s] ∆Videal [m/s] ψR1 [ ] ψR2 [ ] Plesetsk -18.25 198.25 -18.25 211.9 7,926.9 7,995.8 -19.69 199.69 Baiknour -11.85 191.85 -11.85 323.1 7,926.9 7,999.5 -14.11 194.11 CCAFS/KSC -9.35 189.35 - 408.3 7,926.9 8,003.4 -12.23 192.23 VAFB -9.98 189.98 189.98 382.8 7,926.9 8,002.1 -12.68 192.68 Guiana SC -8.23 188.23 -8.23 463.2 7,926.9 8,006.4 -11.52 191.52 Xichang -9.32 189.32 - 409.5 7,926.9 8,003.5 -12.22 192.22 Jiuquan -10.89 190.89 - 351.0 7,926.9 8,000.7 -13.36 193.36 Tanegashima -9.52 189.52 - 401.1 7,926.9 8,003.0 -12.35 192.35 Taiyuan -10.56 190.56 190.56 362.0 7926.9 8,001.2 -13.11 193.11 Satish Shawan -8.44 188.44 - 451.9 7,926.9 8,005.7 -11.64 191.64 MARS -10.41 190.41 - 367.0 7,926.9 8,001.4 -13.00 193.00 Uchinoura -9.61 189.61 - 397.4 7,926.9 8,002.8 -12.42 192.42 Yasny -13.13 193.13 193.13 292.1 7,926.9 7,998.3 -15.17 195.17 Palmachin -9.67 189.67 - 394.8 7,926.9 8,002.7 -12.46 192.46 Imam -10.05 190.05 - 380.0 7,926.9 8,002.0 -12.73 192.73 Vostochny -13.36 193.36 193.36 287.0 7,926.9 7,998.1 -15.37 195.37 Mahia -10.62 190.62 190.62 359.9 7,926.9 8,001.1 -13.16 193.16 PSCA -15.35 195.35 195.35 250.6 7,926.9 7,996.9 -17.08 197.08 Sohae -10.68 190.68 190.68 357.8 7,926.9 8,001.0 -13.20 193.20 Wenchang -8.71 188.71 188.71 438.14 7,926.91 8,005.0 -11.81 191.81

Table B.2: Calculations of the Possibility to Reach the Sun-Synchronous Orbit

Manoeuvres

Elliptical Orbit (Perigee = 200 km and Apogee = 705 km)

s 1 1 uapogee = 2µE( − ) = 7, 361.7 m/s ra rp + ra

Circular Orbit (Radius = 705 km)

rµ u = E = 7, 501.7 m/s circ r

Delta-V needed for manoeuvres:

∆Vman = ucirc − uapogee = 7, 501.7 − 7, 361.7 = 140.0 m/s

81 82 APPENDIX B. POSSIBILITY TO REACH THE ORBITS

B.3 Inclined Orbit

o o o o o Launch Site ψ01 [ ] ψ02 [ ] ψ0 [ ] V0 [m/s] Vf [m/s] ∆Videal [m/s] ψR1 [ ] ψR2 [ ] Plesetsk ------Baiknour 64.43 115.57 64.43 323.1 7,846.7 7,556.5 63.37 116.63 CCAFS/KSC 45.54 134.46 45.54 408.3 7,846.7 7,560.7 43.37 136.63 VAFB 49.57 130.43 - 382.8 7,846.7 7,559.3 47.69 132.31 Guiana SC 38.99 141.01 38.99 463.2 7,846.7 7,563.8 36.26 143.74 Xichang 45.37 134.63 - 409.5 7,846.7 7,560.7 43.19 136.81 Jiuquan 56.13 123.87 - 351.0 7,846.7 7,557.8 54.64 125.36 Tanegashima 46.59 133.41 46.59 401.1 7,846.7 7,560.3 44.50 135.50 Taiyuan 53.62 126.38 - 362.0 7,846.7 7,558.3 52.00 128.00 Satish Shawan 40.16 139.84 40.16 451.9 7,846.7 7,563.1 37.55 142.45 MARS 52.57 127.43 127.43 367.0 7,846.7 7,558.5 50.88 129.12 Uchinoura 47.17 132.83 47.17 397.4 7,846.7 7,560.1 45.12 134.88 Yasny 86.23 93.77 - 292.1 7,846.7 7,555.3 86.08 93.92 Palmachin 47.57 132.43 - 394.8 7,846.7 7,559.9 45.55 134.45 Imam 50.07 129.93 50.07 380.0 7,846.7 7,559.2 48.22 131.78 Vostochny ------Mahia 54.07 125.93 125.93 359.9 7,846.7 7,558.2 52.47 127.53 PSCA ------Sohae 54.53 125.47 - 357.8 7,846.7 7,558.1 52.95 127.05 Wenchang 41.69 138.31 138.31 438.1 7,846.7 7,562.3 39.21 140.79

Table B.3: Calculations of the Possibility to Reach the Inclined Orbit

Manoeuvres

Elliptical Orbit (Perigee = 200 km and Apogee = 415 km)

s 1 1 uapogee = 2µE( − ) = 7, 598.3 m/s ra rp + ra

Circular Orbit (Radius = 415 km)

rµ u = E = 7, 660.2 m/s circ r

Delta-V needed for manoeuvres:

∆Vman = ucirc − uapogee = 7, 660.2 − 7, 598.3 = 61.8 m/s

82 B.4. SEMI-SYNCHRONOUS ORBIT ESEIAAT

B.4 Semi-Synchronous Orbit

o o o o o Launch Site ψ01 [ ] ψ02 [ ] ψ0 [ ] V0 [m/s] Vf [m/s] ∆Videal [m/s] ψR1 [ ] ψR2 [ ] Plesetsk ------Baiknour 46.04 133.96 46.04 323.1 9,898.5 9,668.6 44.71 135.29 CCAFS/KSC 34.71 145.29 - 408.3 9,898.5 9,671.8 32.73 147.27 VAFB 37.40 142.60 - 382.8 9,898.5 9,670.8 35.60 144.40 Guiana SC 30.14 149.86 30.14 463.2 9,898.5 9,674.3 27.76 152.24 Xichang 34.60 145.40 - 409.5 9,898.5 9,671.9 32.60 147.40 Jiuquan 41.49 138.51 138.51 351.0 9,898.5 9,669.6 39.93 140.07 Tanegashima 35.43 144.57 35.43 401.1 9,898.5 9,671.5 33.49 146.51 Taiyuan 39.98 140.02 - 362.0 9,898.5 9,670.0 38.33 141.67 Satish Shawan 30.97 149.03 30.97 451.9 9,898.5 9,673.8 28.68 151.32 MARS 39.32 140.68 140.68 367.0 9,898.5 9,670.2 37.64 142.36 Uchinoura 35.81 144.19 35.81 397.4 9,898.5 9,671.4 33.91 146.09 Yasny 52.77 127.23 - 292.1 9,898.5 9,667.6 51.72 128.28 Palmachin 36.08 143.92 - 394.8 9,898.5 9,671.3 34.19 145.81 Imam 37.73 142.27 37.73 380.0 9,898.5 9,670.7 35.94 144.06 Vostochny 54.13 125.87 54.13 287.0 9,898.5 9,667.5 53.13 126.87 Mahia 40.25 139.75 139.75 359.9 9,898.5 9,669.9 38.62 141.38 PSCA 68.13 111.87 111.87 250.6 9,898.5 9,666.4 67.58 112.42 Sohae 40.53 139.47 - 357.8 9,898.5 9,669.8 38.92 141.08 Wenchang 32.06 147.94 147.94 438.1 9,898.5 9,673.1 29.86 150.14

Table B.4: Calculations of the Possibility to Reach Semi-Synchronous Orbit

Manoeuvres

Elliptical Orbit (Perigee = 200 km and Apogee = 21,000 km)

s 1 1 uapogee = 2µE( − ) = 2, 372.9 m/s ra rp + ra

Circular Orbit (Radius = 21,000 km)

rµ u = E = 3, 815.6 m/s circ r

Delta-V needed for manoeuvres:

∆Vman = ucirc − uapogee = 3, 815.6 − 2, 372.9 = 1443.8 m/s

83 84 APPENDIX B. POSSIBILITY TO REACH THE ORBITS

B.5 Molnyia Orbit

o o o o o Launch Site ψ01 [ ] ψ02 [ ] ψ0 [ ] V0 [m/s] Vf [m/s] ∆Videal [m/s] ψR1 [ ] ψR2 [ ] Plesetsk 79.39 100.61 79.39 211.9 7,881.8 7,673.6 79.10 100.90 Baiknour 40.13 139.87 40.13 323.1 7,881.8 7,677.5 38.29 141.71 CCAFS/KSC 30.66 149.34 - 408.3 7,881.8 7,681.6 28.04 151.96 VAFB 32.95 147.05 - 382.8 7,881.8 7,680.3 30.56 149.44 Guiana SC 26.72 153.28 26.72 463.2 7,881.8 7,684.7 23.63 156.37 Xichang 30.57 149.43 - 409.5 7,881.8 7,681.6 27.94 152.06 Jiuquan 36.39 143.61 143.61 351.0 7,881.8 7,678.7 34.28 145.72 Tanegashima 31.27 148.73 31.27 401.1 7,881.8 7,681.2 28.72 151.28 Taiyuan 35.12 144.88 - 362.0 7,881.8 7,679.2 32.91 147.09 Satish Shawan 27.44 152.56 27.44 451.9 7,881.8 7,684.0 24.45 155.55 MARS 34.57 145.43 145.43 367.0 7,881.8 7,679.5 32.32 147.68 Uchinoura 31.60 148.40 31.60 397.4 7,881.8 7,681.0 29.08 150.92 Yasny 45.48 134.52 134.52 292.1 7,881.8 7,676.3 43.95 136.05 Palmachin 31.83 148.17 - 394.8 7,881.8 7,680.9 29.33 150.67 Imam 33.23 146.77 33.23 380.0 7,881.8 7,680.1 30.85 149.15 Vostochny 46.52 133.48 46.52 287.0 7,881.8 7,676.1 45.05 134.95 Mahia 35.35 144.65 144.65 359.9 7,881.8 7,679.1 33.16 146.84 PSCA 56.21 123.79 123.79 250.6 7,881.8 7,674.8 55.17 124.83 Sohae 35.59 144.41 - 357.8 7,881.8 7,679.0 33.42 146.58 Wenchang 28.38 151.62 151.62 438.1 7,881.8 7,683.2 25.50 154.50

Table B.5: Calculations of the Possibility to Reach the Molniya Orbit

Manoeuvres

Elliptical Orbit (Perigee = 200 km and Apogee = 540 km)

s 1 1 uapogee,1 = 2µE( − ) = 7, 494.4 m/s ra rp + ra

Elliptical Orbit (Perigee = 540 km and Apogee = 39,300 km)

s 1 1 uperigee,2 = 2µE( − ) = 10, 003.9 m/s rp rp + ra

Delta-V needed for manoeuvres:

∆Vman = uperigee,2 − uapogee,1 = 10, 003.9 − 7, 494.4 = 2, 509.5 m/s

84 Appendix C

Results of Optimal Staging

*Note:

N2O4 represents N2O4/UDMH RP − 1 represents LOX/RP − 1

LH2 represents LOX/LH2

85 86 APPENDIX C. RESULTS OF OPTIMAL STAGING

C.1 Polar Orbit

Payload LOX/RP-1 & LOX/RP-1 LOX/LH2 & LOX/LH2

Take-Off Mass [kg] 108,000 24,500 Thrust (x 1.5) [kN] 1,589 361 1,000 kg Thrust (x 1.2) [kN] 1,271 288 Beta Delta IV

Take-Off Mass [kg] 54,000 12,250 Thrust (x 1.5) [kN] 795 180 500 kg Thrust (x 1.2) [kN] 636 144 Alpha

Take-Off Mass [kg] 10,800 2,450 Thrust (x 1.5) [kN] 159 36 100 kg Thrust (x 1.2) [kN] 127 29 Electron

Take-Off Mass [kg] 1,080 245 Thrust (x 1.5) [kN] 16 4 10 kg Thrust (x 1.2) [kN] 13 3 Vector-R

Table C.1: Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Polar Orbit

Payload 3x N2O4 / UDMH 2x LOX / RP-1 & N2O4 / UDMH 3x LOX / RP-1

Take-Off Mass [kg] 66,000 60,400 56,000 Thrust (x 1.5) [kN] 971 889 824 1,000 kg Thrust (x 1.2) [kN] 777 711 659 Rockot Soyuz 2.1v Zenit Strela

Take-Off Mass [kg] 33,000 30,200 28,000 500 kg Thrust (x 1.5) [kN] 486 444 412 Thrust (x 1.2) [kN] 388 356 330

Take-Off Mass [kg] 6,600 6,040 5,600 100 kg Thrust (x 1.5) [kN] 97 89 82 Thrust (x 1.2) [kN] 78 71 66

Take-Off Mass [kg] 660 604 560 10 kg Thrust (x 1.5) [kN] 10 9 8 Thrust (x 1.2) [kN] 8 7 7

Table C.2: Optimal Three Stage Launch Vehicles according to the Propellant Type to Reach Polar Orbit.

86 C.2. SUN-SYNCHRONOUS ORBIT ESEIAAT

Payload 4x Solid Solid & N2O4 & 2x Solid 4x N2O4 3x Solid & N2O4 5x Solid

Take-Off Mass [kg] 98,600 80,400 56,000 80,400 87,000 Thrust (x 1.5) [kN] 1,451 1,183 824 1,183 1,280 1,000 kg Thrust (x 1.2) [kN] 1,161 946 659 946 1,024 Vega PSLV-CA Protom M SSLV Minotaur IV

Take-Off Mass [kg] 49,300 40,200 28,000 40,200 43,500 Thrust (x 1.5) [kN] 725 592 412 592 640 500 kg Thrust (x 1.2) [kN] 580 473 330 473 512 Minotaur C

Take-Off Mass [kg] 9,860 8,040 5,600 8,040 8,700 Thrust (x 1.5) [kN] 145 118 82 118 128 100 kg Thrust (x 1.2) [kN] 116 95 66 95 102 Minotaur I

Take-Off Mass [kg] 986 804 560 804 870 10 kg Thrust (x 1.5) [kN] 15 12 8 12 13 Thrust (x 1.2) [kN] 12 9 7 9 10

Table C.3: Optimal Four and Five Stage Launch Vehicles according to the Propellant Type∗ to Reach Polar Orbit.

C.2 Sun-Synchronous Orbit

Payload LOX/RP-1 & LOX/RP-1 LOX/LH2 & LOX/LH2

Take-Off Mass [kg] 91,000 22,300 Thrust (x 1.5) [kN] 1,339 328 1,000 kg Thrust (x 1.2) [kN] 1,071 263 Beta Delta IV

Take-Off Mass [kg] 45,500 11,150 Thrust (x 1.5) [kN] 670 164 500 kg Thrust (x 1.2) [kN] 536 131 Alpha

Take-Off Mass [kg] 9,100 2,230 Thrust (x 1.5) [kN] 134 33 100 kg Thrust (x 1.2) [kN] 107 26 Electron Vector-H

Take-Off Mass [kg] 910 223 Thrust (x 1.5) [kN] 13 3 10 kg Thrust (x 1.2) [kN] 11 3 Vector-R

Table C.4: Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Sun-Synchronous Orbit

87 88 APPENDIX C. RESULTS OF OPTIMAL STAGING

Payload 4x Solid 5x Solid

Take-Off Mass [kg] 87,000 77,100 Thrust (x 1.5) [kN] 1,280 1,135 1,000 kg Thrust (x 1.2) [kN] 1,024 908 Minotaur IV

Take-Off Mass [kg] 43,500 38,550 Thrust (x 1.5) [kN] 640 567 500 kg Thrust (x 1.2) [kN] 512 454 Minotaur C

Take-Off Mass [kg] 8,700 7,710 Thrust (x 1.5) [kN] 128 113 100 kg Thrust (x 1.2) [kN] 102 91 Minotaur I

Take-Off Mass [kg] 870 771 10 kg Thrust (x 1.5) [kN] 13 11 Thrust (x 1.2) [kN] 10 9

Table C.5: Optimal Four and Five Stage Launch Vehicles according to the Propellant Type to Reach Sun-Synchronous Orbit.

88 C.3. INCLINED ORBIT ESEIAAT

C.3 Inclined Orbit

Payload LOX/RP-1 & LOX/RP-1 LOX/LH2 & LOX/LH2

Take-Off Mass [kg] 65,000 18,400 Thrust (x 1.5) [kN] 956 271 1,000 kg Thrust (x 1.2) [kN] 765 217 Beta Delta IV

Take-Off Mass [kg] 32,500 9,200 Thrust (x 1.5) [kN] 478 135 500 kg Thrust (x 1.2) [kN] 383 108 Alpha

Take-Off Mass [kg] 6,500 1,840 Thrust (x 1.5) [kN] 96 27 100 kg Thrust (x 1.2) [kN] 77 22 Vector-R

Take-Off Mass [kg] 650 184 10 kg Thrust (x 1.5) [kN] 10 3 Thrust (x 1.2) [kN] 8 2

Table C.6: Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Inclined Orbit

Payload 3x N2O4 2x RP-1 & N2O4 3x RP-1 2x N2O4 & Solid RP-1 & 2x Solid 3x LH2

Take-Off Mass [kg] 46,000 42,300 39,300 55,500 70,600 15,000 Thrust (x 1.5) [kN] 677 622 578 817 1,039 221 1,000 kg Thrust (x 1.2) [kN] 542 498 463 653 831 177 Strela Soyuz 2.1v LM7 Antares LM5

Take-Off Mass [kg] 23,000 21,150 19,650 27,750 35,300 7,500 Thrust (x 1.5) [kN] 338 311 289 408 519 110 500 kg Thrust (x 1.2) [kN] 271 249 231 327 416 88 Safir

Take-Off Mass [kg] 4,600 4,230 3,930 5,550 7,060 1,500 100 kg Thrust (x 1.5) [kN] 68 62 58 82 104 22 Thrust (x 1.2) [kN] 54 50 46 65 83 18

Take-Off Mass [kg] 460 423 393 555 706 150 10 kg Thrust (x 1.5) [kN] 7 6 6 8 10 2 Thrust (x 1.2) [kN] 5 5 5 7 8 2

Table C.7: Optimal Two Stage Launch Vehicles according to the Propellant Type∗ to Reach Inclined Orbit

89 90 APPENDIX C. RESULTS OF OPTIMAL STAGING

Payload 4x Solid Solid & N2O4 & 2x Solid 4x N2O4 3x Solid & N2O4 5x Solid

Take-Off Mass [kg] 67,300 59,400 40,000 60,000 60,500 Thrust (x 1.5) [kN] 990 874 589 883 890 1,000 kg Thrust (x 1.2) [kN] 792 699 471 706 712 Minotar C PSLV-CA Protom M SSLV Minotaur IV

Take-Off Mass [kg] 33,650 29,700 20,000 30,000 30,250 Thrust (x 1.5) [kN] 495 437 294 441 445 500 kg Thrust (x 1.2) [kN] 396 350 235 353 356 Minotaur I

Take-Off Mass [kg] 6,730 5,940 4,000 6,000 6,050 100 kg Thrust (x 1.5) [kN] 99 87 59 88 89 Thrust (x 1.2) [kN] 79 70 47 71 71

Take-Off Mass [kg] 673 594 400 600 605 10 kg Thrust (x 1.5) [kN] 10 9 6 9 9 Thrust (x 1.2) [kN] 8 7 5 7 7

Table C.8: Optimal Four and Five Stage Launch Vehicles according to the Propellant Type∗ to Reach Inclined Orbit.

C.4 Semi-Synchronous Orbit

Payload LOX/RP-1 & LOX/RP-1 LOX/LH2 & LOX/LH2 N2O4/UDMH & N2O4/UDMH

Take-Off Mass [kg] 1,700,000 76,000 3,200,000 Thrust (x 1.5) [kN] 25,016 1,118 47,088 1,000 kg Thrust (x 1.2) [kN] 20,012 895 37,670 HII-A

Take-Off Mass [kg] 850,000 38,000 1,600,000 500 kg Thrust (x 1.5) [kN] 12,508 559 23,544 Thrust (x 1.2) [kN] 10,006 447 18,835

Take-Off Mass [kg] 170,000 7,600 320,000 Thrust (x 1.5) [kN] 2,502 112 4,709 100 kg Thrust (x 1.2) [kN] 2,001 89 3,767 Angara 1.2

Take-Off Mass [kg] 17,000 760 32,000 Thrust (x 1.5) [kN] 250 11 471 10 kg Thrust (x 1.2) [kN] 200 9 377 LM2D

Table C.9: Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Semi-Synchronous Orbit

90 C.4. SEMI-SYNCHRONOUS ORBIT ESEIAAT

Payload 3x N2O4 2x RP-1 & N2O4 3x RP-1 2x N2O4 & Solid RP-1 & 2x Solid 3x Solid 3x LH2

Take-Off Mass [kg] 263,000 233,000 208,000 362,000 535,000 692,000 46,000 Thrust (x 1.5) [kN] 3,870 3,429 3,061 5,327 7,873 10,183 677 1,000 kg Thrust (x 1.2) [kN] 3,096 2,743 2,449 4,261 6,298 8,146 542 Proton Medium Soyuz 2.1a/b LM7 LM5

Take-Off Mass [kg] 131,500 116,500 104,000 181,000 267,500 346,000 23,000 Thrust (x 1.5) [kN] 1,935 1,714 1,530 2,663 3,936 5,091 338 500 kg Thrust (x 1.2) [kN] 1,548 1,371 1,224 2,131 3,149 4,073 271 LM4B Soyuz 2.1v LM2C LM4C

Take-Off Mass [kg] 26,300 23,300 20,800 36,200 53,500 69,200 4,600 Thrust (x 1.5) [kN] 387 343 306 533 787 1018 68 100 kg Thrust (x 1.2) [kN] 310 274 245 426 630 815 54 Strela Antares

Take-Off Mass [kg] 2,630 2,330 2,080 3,620 5,350 6,920 460 Thrust (x 1.5) [kN] 39 34 31 53 79 102 7 10 kg Thrust (x 1.2) [kN] 31 27 24 43 63 81 5 Safir Zhuque-1

Table C.10: Optimal Three Stage Launch Vehicles according to the Propellant Type∗ to Reach Semi-Synchronous Orbit.

Payload 4x Solid Solid & N2O4 & 2x Solid 4x N2O4 3x Solid & N2O4 5x Solid

Take-Off Mass [kg] 402,000 310,000 189,000 315,000 319,000 Thrust (x 1.5) [kN] 5,915 4,562 2,781 4,635 4,694 1,000 kg Thrust (x 1.2) [kN] 4,732 3,649 2,225 3,708 3,755 Protom M

Take-Off Mass [kg] 201,000 155,000 94,500 157,500 159,500 Thrust (x 1.5) [kN] 2,958 2,281 1,391 2,318 2,347 500 kg Thrust (x 1.2) [kN] 2,366 1,825 1,112 1,854 1,878 PSLV-CA Vega C

Take-Off Mass [kg] 40,200 31,000 18,900 31,500 31,900 Thrust (x 1.5) [kN] 592 456 278 464 469 100 kg Thrust (x 1.2) [kN] 473 365 222 371 376 Hyperbola-I Kuaizhou 1 Minotaur IV

Take-Off Mass [kg] 4,020 3,100 1,890 3,150 3,190 10 kg Thrust (x 1.5) [kN] 59 46 28 46 47 Thrust (x 1.2) [kN] 47 36 22 37 38

Table C.11: Optimal Four and Five Stage Launch Vehicles according to the Propellant Type∗ to Reach Semi-Synchronous Orbit.

Payload N2O4 & LH2 2x RP-1 2x N2O4 2x RP-1 & N2O4 Solid & N2O4 & LH2

Take-Off Mass [kg] 140,000 357,000 413,000 211,000 120,000 Thrust (x 1.5) [kN] 2,060 5,253 6,077 3,105 1,766 1,000 kg Thrust (x 1.2) [kN] 1,648 4,203 4,862 2,484 1,413 LVM3 Angara 3 LM2F Soyuz 2.1a/b GLSV

Table C.12: Optimal Launch Vehicles with Boosters Configuration according to the Propellant Type∗ to Reach Semi-Synchronous Orbit.

91 92 APPENDIX C. RESULTS OF OPTIMAL STAGING

C.5 Molnyia Orbit

Payload LOX/RP-1 & LOX/RP-1 LOX/LH2 & LOX/LH2 N2O4/UDMH & N2O4/UDMH

Take-Off Mass [kg] 458,000 49,000 640,000 Thrust (x 1.5) [kN] 6,739 721 9,418 1,000 kg Thrust (x 1.2) [kN] 5,392 577 7,534 HII-A

Take-Off Mass [kg] 229,000 24,500 320,000 500 kg Thrust (x 1.5) [kN] 3,370 361 4,709 Thrust (x 1.2) [kN] 2,696 288 3,767

Take-Off Mass [kg] 45,800 4,900 64,000 Thrust (x 1.5) [kN] 674 72 942 100 kg Thrust (x 1.2) [kN] 539 58 753 Angara 1.2 LM2D

Take-Off Mass [kg] 4,580 490 6,400 Thrust (x 1.5) [kN] 67 7 94 10 kg Thrust (x 1.2) [kN] 54 6 75 Vector-R

Table C.13: Optimal Two Stage Launch Vehicles according to the Propellant Type to Reach Molnyia Orbit

Payload 3x N2O4 2x RP-1 & N2O4 3x RP-1 2x N2O4 & Solid RP-1 & 2x Solid 3x Solid 3x LH2

Take-Off Mass [kg] 155,000 139,500 126,000 205,000 284,500 355,000 33,300 Thrust (x 1.5) [kN] 2,281 2,053 1,854 3,017 4,186 5,224 490 1,000 kg Thrust (x 1.2) [kN] 1,825 1,642 1,483 2,413 3,349 4,179 392 Dnepr Soyuz 2.1v Zenit LM2C LM5

Take-Off Mass [kg] 77,500 69,750 63,000 102,500 142,250 177,500 16,650 Thrust (x 1.5) [kN] 1,140 1,026 927 1,508 2,093 2,612 245 500 kg Thrust (x 1.2) [kN] 912 821 742 1,207 1,675 2,090 196 Rockot Antares Strela

Take-Off Mass [kg] 15,500 13,950 12,600 20,500 28,450 35,500 3,330 Thrust (x 1.5) [kN] 228 205 185 302 419 522 49 100 kg Thrust (x 1.2) [kN] 182 164 148 241 335 418 39 Safir

Take-Off Mass [kg] 1,550 1,395 1,260 2,050 2,845 3,550 333 Thrust (x 1.5) [kN] 23 21 19 30 42 52 5 10 kg Thrust (x 1.2) [kN] 18 16 15 24 33 42 4 Zhuque-1

Table C.14: Optimal Three Stage Launch Vehicles according to the Propellant Type∗ to Reach Molnyia Orbit.

92 C.5. MOLNYIA ORBIT ESEIAAT

Payload 4x Solid Solid & N2O4 & 2x Solid 4x N2O4 3x Solid & N2O4 5x Solid

Take-Off Mass [kg] 238,000 187,000 121,000 205,000 197,500 Thrust (x 1.5) [kN] 3,502 2,752 1,781 3,017 2,906 1,000 kg Thrust (x 1.2) [kN] 2,802 2,201 1,424 2,413 2,325 PSLV-CA Protom M Vega C

Take-Off Mass [kg] 119,000 93,500 60,500 102,500 98,750 Thrust (x 1.5) [kN] 1,751 1,376 890 1,508 1,453 500 kg Thrust (x 1.2) [kN] 1,401 1,101 712 1,207 1,162 Vega SSLV

Take-Off Mass [kg] 23,800 18,700 12,100 20,500 19,750 Thrust (x 1.5) [kN] 350 275 178 302 291 100 kg Thrust (x 1.2) [kN] 280 220 142 241 232 Minotaur I Kuaizhou 1 Minotaur IV

Take-Off Mass [kg] 2,380 1,870 1,210 2,050 1,975 10 kg Thrust (x 1.5) [kN] 35 28 18 30 29 Thrust (x 1.2) [kN] 28 22 14 24 23

Table C.15: Optimal Four and Five Stage Launch Vehicles according to the Propellant Type∗ to Reach Molnyia Orbit.

Payload 2x RP-1 2x N2O4 2x RP-1 & N2O4 Solid & N2O4 & LH2

Take-Off Mass [kg] 203,000 231,000 133,000 76,000 Thrust (x 1.5) [kN] 2,987 3,399 1,957 1,118 1,000 kg Thrust (x 1.2) [kN] 2,390 2,719 1,566 895 Angara 3 LM2F Soyuz 2.1a/b GLSV

Table C.16: Optimal Launch Vehicles with Boosters Configuration according to the Propellant Type∗ to Reach Molnyia Orbit.

93

Appendix D

Optimisation Script

%%% Generic Script %%%

% Constant g0 = 9 . 8 0 6 6 ; beta= sqrt(2)/2; e r r o r =1e −4;

% Input Data DeltaV_id = ;%Ideal DeltaV Drag = 200;%Draf Losses Gravity = 1550;%Gravity Losses Man = ;%Manoeuvres Steering = 4;%Steering Losses(Percentage) DeltaV = (DeltaV_id + Drag + Gravity)/(1 − Steering/100) + Man; Payload = ;

% LV Performance N = ;%Number of Stages

Isp = [ ] ;%Specific Impulse

a0g0(1) = 1.25;

f o r i=2:N a0g0(i) = 0.35; end

95 96 APPENDIX D. OPTIMISATION SCRIPT

TW( 1 ) = 8 0 ; f o r i=2:N TW( i ) = 3 0 ; end

kt = [ ] ;%Tank Contribution

m0(N+1) = Payload;

f o r i=1:N c(i) = Isp(i) ∗g0 ; MR_guess(i) = 0.1; end

hyp = 1 ;

while hyp ~= 0 f o r i=1:N ks(i) = (kt(i) ∗(1 − MR_guess(i)) + a0g0(i)/TW(i)) /((1 + kt ( i ) ) ∗(1 − MR_guess(i)) + a0g0(i)/TW(i)); Ar_mu0( i ) = beta/(c(i) ∗(beta − ks ( i ) ) ) ; Ar(i) = 1/c(i); end

mu0 = max(Ar_mu0) ; mu = mu0 ; mu_min = Ar( i ) ; cond1 = 1 ;

while cond1 ~= 0 g = DeltaV ; f o r i=1:N g = g + c ( i ) ∗ l o g (mu∗c ( i ) ∗ ks ( i ) /(mu∗c ( i ) − 1) ) ; end

i f abs(g) < error f o r i=N: −1:1 MR( i ) = mu∗c ( i ) ∗ ks ( i ) /(mu∗c ( i ) − 1) ; deltaV ( i ) = −c ( i ) ∗ l o g (MR( i ) ) ; ratio(i) = deltaV(i)/DeltaV; lambda(i) = (MR(i) − ks ( i ) ) /(1 − ks ( i ) ) ; m0(i) = m0(i + 1)/lambda(i);

96 ESEIAAT

mp( i ) = (1 − MR( i ) ) ∗m0( i ) ; ms(i) = ks(i)/(1 − ks ( i ) ) ∗mp( i ) ; end cond1 = 0 ; hyp = 0 ; f o r i=1:N Hyp( i ) = 1 ; i f abs(MR_guess( i ) − MR( i ) ) > error MR_guess(i) = MR(i); e l s e Hyp( i ) = 0 ; end hyp = hyp + Hyp(i); end e l s e Dg = 0 ; f o r i=1:N Dg = Dg − c ( i ) /(mu∗(mu∗c ( i ) − 1) ) ; end Dmu = − g/Dg ; alpha = 1 ; cond2 = 1 ;

while cond2 ~= 0 mu_new = mu + alpha ∗Dmu; i f mu_new>mu_min mu = mu_new; cond2 = 0 ; e l s e alpha = alpha/2; end end end end end

97

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103 2 ESEIAAT

I declare that,

the work in this Master Thesis / Degree Thesis (choose one) is completely my own work,

no part of this Master Thesis / Degree Thesis (choose one) is taken from other people’s work without giving them credit,

all references have been clearly cited,

I’m authorised to make use of the company’s / research group (choose one) related information I’m providing in this document (select when it applies).

I understand that an infringement of this declaration leaves me subject to the foreseen disciplinary actions by The Universitat Politècnica de Catalunya - BarcelonaTECH.

Robert Arcaleanu ______30/06/2020 Student Name Signature Date

Title of the Thesis : Study of the feasibility of a “Rocket launching Consultancy” through the analysis of the propulsion’s systems requirements to reach LEO and MEO orbits with payloads up to 1,000 kg